ES2985102T3 - Fluidic propulsion system and thrust and lift generator for aerial vehicles - Google Patents

Abstract

A vehicle comprises a main body and a gas generator producing a gas stream. At least one front duct and one rear duct are fluidly coupled to the generator. A first and a second front ejector are fluidly coupled to the at least one front duct. At least one rear ejector is fluidly coupled to the at least one rear duct. The front ejectors respectively include an outlet structure from which gas exits the at least one front duct. The at least one rear ejector includes an outlet structure from which gas exits the at least one rear duct. A first and a second primary aerodynamic element have leading edges respectively located directly downstream of the first and the second front ejector. At least one secondary aerodynamic element has a leading edge located directly downstream of the outlet structure of the at least one rear ejector. (Automatic translation with Google Translate, without legal value)

Description

DESCRIPTION

Fluidic propulsion system and thrust and lift generator for aerial vehicles

Copyright Notice

This disclosure is protected under United States and International copyright laws. © 2016 Jetoptera. All rights reserved. A portion of the disclosure in this patent document contains material that is subject to copyright protection. The copyright owner has no objection to facsimile reproduction by anyone of the patent document or patent disclosure as it appears in the patent file or records of the Patent and Trademark Office, but otherwise all copyright whatever is reserved.

Priority claim

This application claims priority to U.S. Provisional Application No. 62/213,465, filed Sept. 2, 2015.

Background

Aircraft that can hover, take off, and land vertically are commonly referred to as vertical takeoff and landing (VTOL) aircraft. This classification includes fixed-wing aircraft as well as helicopters and aircraft with powered tiltrotors. Some VTOL aircraft can also operate in other modes, such as short takeoff and landing (STOL). VTOL is a subset of V/STOL (vertical and/or short takeoff and landing).

For illustrative purposes, an example of a current aircraft that has VTOL capability is the F-35 Lightning. Conventional methods of vectoring vertical lift airflow include the use of nozzles that can be rotated in a single direction along with the use of two sets of flat flap vanes arranged 90 degrees to each other and located on the outboard nozzle. The F-35 Lightning propulsion system similarly provides vertical lift using a combination of vectored thrust from the turbine engine and a vertically oriented lift fan. The lift fan is located aft of the cockpit in a compartment with upper and lower clamshell doors. The engine is exhausted through a three-bearing rotating nozzle that can deflect thrust from the horizontal to just forward of the vertical. Roll control ducts extend outward on each wing and are supplied with air thrust from the engine fan. Pitch control is effected by a lift fan/engine thrust split. Yaw control is through the yaw motion of the engine-mounted rotating nozzle. Roll control is provided by differentially opening and closing the openings at the ends of the two roll control ducts. The lift fan has a "D" shaped telescoping nozzle to provide thrust deflection in the forward and aft directions. The D-shaped nozzle has fixed vanes at the exit opening.

The design of an aircraft or drone most generally consists of its propulsion elements and the fuselage into which those elements are integrated. Classically, the propulsion device in aircraft may be a jet engine, turbofan, turboprop or turboshaft, piston engine, or an electric motor equipped with a propeller. The propulsion system (thruster) in small unmanned aerial vehicles (UAVs) is conventionally a piston engine or electric motor providing power through a shaft to one or more propellers. The propulsion system for a larger aircraft, whether manned or unmanned, is traditionally a jet engine or turboprop. The propulsion system is usually attached to the fuselage or body or wings of the aircraft through pylons or struts capable of transmitting force to the aircraft and maintaining loads. The emerging mixed jet (jet outflow) of air and gases is what propels the aircraft in the opposite direction to the jet outflow flow.

Conventionally, the airflow outflow from a large propeller is not used for lift purposes in horizontal flight, and thus no significant amount of kinetic energy is used for the benefit of the aircraft, unless it is rotated as in some of the currently existing applications (namely the Bell Boeing V-22 Osprey). Instead, lift on most existing aircraft is created by the wings and tail. Furthermore, even in those particular VTOL applications (e.g. takeoff via transition to horizontal flight) found on the Osprey, the lift caused by the propeller itself is minimal during horizontal flight, and most of the lift force is nevertheless from the wings.

The current state of the art for creating lift on an aircraft is to generate a high-speed airflow over the wing and wing elements, which are generally lifting bodies. Lifting bodies are characterized by a chord line extending mainly in the axial direction, from a leading edge to a trailing edge of the lifting body. Based on the angle of attack formed between the incident airflow and the chord line, and according to the principles of lifting body lift generation, lower pressure air flows over the suction (upper) side and, conversely, by Bernoulli's law, moves at higher velocities than the lower side (pressure side). The lower the aircraft airspeed, the lower the lift force and the larger the wing surface area or higher angles of incidence are required, even for takeoff.

Large UAVs make no exception to this rule. Lift is generated by designing a wing lifting body with the appropriate angle of attack, chord, span and bank line. Flaps, slots and many other devices are other conventional tools used to maximize lift by increasing the coefficient of lift and wing surface area, but they will generate lift corresponding to the aircraft's airspeed. (Increasing the area (S) and coefficient of lift (CL-) allow a similar amount of lift to be generated at a lower aircraft airspeed (V0) according to the formula L = 1/2 pV2SCl, but at the cost of higher drag and weight.) These current techniques also perform poorly with a significant drop in efficiency in high crosswind conditions.

Although smaller UAVs arguably use propeller-generated thrust to lift the vehicle, current technology relies strictly on control of electric motor speeds, and the smallest UAV may or may not have the ability to spin the motors to generate thrust and lift, or transition to horizontal flight by pitching the propellers. In addition, smaller UAVs that use these propulsion elements suffer from inefficiencies related to batteries, power density, and large propellers, which may be efficient in hovering but inefficient in horizontal flight and create difficulty and danger when operating due to the fast-moving tip of the blades. Most current quadcopters and other electrically powered aerial vehicles are only capable of very short flight periods and cannot efficiently lift or carry large payloads, as the weight of the electric motor and battery system can already exceed 70% of the vehicle weight at any given time in flight. A similar vehicle using jet fuel or any other hydrocarbon fuel typically used in transportation will carry more usable fuel by at least an order of magnitude. This can be explained by the much higher energy density of hydrocarbon fuel compared to battery systems (by at least an order of magnitude), as well as the lower weight to total vehicle weight ratio of a hydrocarbon fuel-based system.

Therefore, there is a need for improved efficiency, enhanced capabilities and other technological advances in aircraft, particularly for UAVs and certain manned aerial vehicles.

Document CN 103 419 933 discloses a vertical takeoff and landing aircraft with forward wings and rear wings configured to provide high lift.

Brief description of the drawing figures

FIGS. 1A - 1C illustrate some of the differences in structure, forces and controls between a conventional electric quadcopter and an example that does not fall within the scope of the claims.

FIG. 2A is a top view of a conventional wing and aircraft structure; FIG. 2B is a front view of a conventional wing and aircraft structure.

FIG. 3 is a cross section of an example not falling within the scope of the claims depicting only the upper half of an ejector and showing velocity and temperature profiles within the internal flow.

FIG 4 is an example not falling within the scope of the claims depicting a thruster/ejector positioned in front of a lifting body.

FIG. 5 is another example not within the scope of the claims where the propeller/ejector is placed in front of a control surface as part of another wing lifting body.

FIGS. 6A - 6C illustrate the example shown in FIG. 5 from different points of view.

FIG 7A is another example not within the scope of the claims that uses a jet outflow and a lifting body in its wake to push the aircraft forward and generate lift, replacing the engine on the wing.

FIG. 7B is the front view of the example shown in FIG. 7A.

FIG 7C is another example not falling within the scope of the claims that features tandem wings.

FIG. 8A is a side view of an embodiment of the present invention, showing the tandem thrust/lift generating system where the forward thrust-augmenting ejectors are producing thrust with a canard wing and the rear thrust-augmenting ejectors are producing thrust and lift in the rear region.

FIG. 8B is the perspective view of the present invention shown in FIG. 8A.

FIG. 9 is a perspective view of the present invention shown in FIGS. 8A and 8B and features the aircraft tail arrangements and gas generator assembly.

FIGS. 10A - 10E show variations of lift coefficient at a constant airspeed of a lifting body as a function of incidence angle showing the stall angle of attack.

FIGS. 11A - 11B show an improvement in the stall margin.

FIGS. 12A - 12C are another example not within the scope of the claims that presents the ejector component of the propeller in a position relative to the wing.

FIGS. 13A - 13C illustrate how the present invention can control the pitch, roll and yaw of the aircraft using the thrust augmentation ejectors in conjunction with thin lifting bodies placed in the wake of the ejectors.

FIG 14 is an example not falling within the scope of the claims with fin-like elements to the diffuser walls of a Coanda ejector that is divided into two halves.

FIGS. 15A - 15C illustrate 3D features of an example not falling within the scope of the claims, from different viewpoints.

FIG 16A shows another example not falling within the scope of the claims for improving performance and stall clearance.

FIGS. 16B - 16D illustrate the example shown in FIG. 16A from different viewpoints.

FIGS. 17A-17C illustrate yet another example that does not fall within the scope of the claims.

FIGS. 18A - 18C show typical conventional arrangements for Coanda type ejectors.

FIG 18D is an example not within the scope of the claims that depicts a circular Coanda ejector with simple primary nozzle elements.

FIGS. 19A-19D depict different examples not falling within the scope of the claims that feature improved performing primary nozzles.

FIG. 19E illustrates flow over a delta wing obstruction positioned within the primary nozzle at its center. FIG. 20 explains the thermodynamics of an example not within the scope of the claims. FIG. 21 is yet another example not within the scope of the claims and features for improving flow separation delay.

FIGS. 22A to 22F illustrate various 3D features and examples that do not fall within the scope of the claims.

FIG. 23 illustrates certain features according to an example not falling within the scope of the claims. FIG. 24 shows a Coanda type ejector applied to a VTOL aircraft only.

FIG. 25 shows an alternative ejector arrangement as another example that does not fall within the scope of the claims.

FIG. 26A shows a high bypass turbofan.

FIG. 26B shows a modified turbofan serving as a gas generator.

FIG. 27A is an embodiment of the present invention showing the purge network and ducts.

FIG. 27B is another embodiment of a purge and duct network.

FIG. 27C is yet another embodiment of a purge and duct network showing the controller and sensors. FIG. 27D is yet another embodiment of a purge and duct network showing the controller and identified sensors.

FIGS. 28A - 28E are possible forms of propellants.

FIG. 29 is a possible arrangement of the propulsion system in takeoff or hover in an embodiment of the present invention.

FIGS. 30A - 30B illustrate the thermodynamic cycles of a jet engine.

FIG. 31 is an embodiment of the present invention.

Detailed description

This application is intended to describe one or more embodiments of the present invention. It is to be understood that the use of absolute terms, such as "should," "will," and the like, as well as specific quantities, is to be construed as applicable to one or more such embodiments, but not necessarily to all such embodiments. As such, embodiments of the invention may omit or include a modification, one or more features or functionality described in the context of such absolute terms. Furthermore, the headings in this application are for reference purposes only and shall not in any way affect the meaning or interpretation of the present invention. The present inventions disclosed in this application, whether independently and working together, enable a UAV to perform the maneuvers of an electric UAV without the use of large propellers or fans while also maximizing the vehicle's range, reach, and payload to total weight ratio. Electric UAVs such as a quadcopter can hover, take off and land vertically as such, perform loops, etc. simply by controlling the rotational speed of the propellers attached to it. The present invention eliminates the need for propellers or large fans and replaces the propeller rotational speed control logic with primarily fluidic control of rotary thrust augmentation ejectors supplied with a motive fluid from an onboard gas generator. Non-electric UAVs employing jet engines typically do not operate at low speeds or efficiently and are limited in their maneuverability compared to electric UAVs. FIGS. 1A through 1C illustrate some of the differences in structure, forces, and rotational speeds between a conventional electric quadcopter and a fluidic quadcopter.

The present invention introduces a vehicle according to claim 1. Preferred embodiments are covered in the dependent claims.

Propulsion device and thrust system

FIGS. 2A and 2B depict a conventional aircraft with wing-mounted thrust-producing engines, which generate acceleration and speed of the aircraft, resulting in the generation of lift on the wings; the function of the engine is to create the thrust and the jet outflow from the engine is not used for additional lift generation, but is lost to the environment. The jet outflow has a velocity greater than that of the aircraft, and as such, the lift generated by the wing is a function of the aircraft's airspeed and not the local engine jet outflow velocity, which is the subject of the current application.

An example not within the scope of the claims includes a thruster which utilizes fluids for entrainment and acceleration of ambient air and supplies a high velocity jet outflow of a mixture of high pressure gas (supplied to the thruster from a gas generator) and entrained ambient air in a manner designed directly toward a lifting body positioned just behind the thruster, in the wake of the propellant jet, and in a symmetrical or non-symmetrical manner.

FIG. 3 illustrates a cross section of only the upper half of ejector 200. Plenum chamber 211 is supplied with air warmer than ambient. The pressurized motive gas stream 600, which may be produced by, for example but not limited to, a fossil fuel combustion engine or an electrically powered compressor, is communicated through passages with primary nozzles 203 to the inner side of the ejector. The primary nozzles accelerate the motive fluid 600 to the velocity required for ejector performance, by design of the primary nozzles 203. The primary (motive) fluid 600 emerges at high velocity over Coanda surface 204 as a wall jet, entraining ambient air 1 which may be at rest or approaching the ejector at non-zero velocity from the left of the figure. The mixture of stream 600 and ambient 1 moves purely axially in throat section 225 of the ejector. Through diffusion in the diffuser 210, the mixing and smoothing process continues, so that the temperature (750) and velocity profiles in the axial direction (700) no longer have high and low values as they do in the throat section 225, but become more uniform at the ejector exit. As the mixture of 1 and 600 approaches the exit plane, the temperature and velocity profiles are nearly uniform; in particular, the temperature of the mixture is low enough to be directed toward a lifting body such as a wing or control surface.

4, another example not falling within the scope of the claims is illustrated, with the thruster/ejector 200 positioned ahead of a lifting body 100 and generating a lift force 400. The local flow over the lifting body 100 is at a higher velocity than the aircraft velocity, due to the higher velocity 300 of the exit jet outflow from the thruster 200 compared to the aircraft airspeed 500. The thruster vigorously mixes a warmer motive stream provided by the gas generator with the incoming cold ambient stream of air at a high drag velocity. The mixture is homogeneous enough to reduce the hot motive stream 600 from the ejector temperature to a mixture temperature profile 750 that will not impact the lifting bodies 100 or 150 mechanically or structurally. The velocity profile of the outflow jet exiting the booster 200 is such that it will allow more lift 400 to be generated by the lifting body 100 due to higher local velocities. Additional control surfaces may be implemented on the lifting body 100, such as the lifting surface 150 shown here. By changing the angle of such surfaces 150, the attitude of the aircraft may be changed rapidly with little effort given the higher local velocity of the outflow jet 300.

FIG. 5 illustrates that the thruster/ejector 200 may also be positioned forward of a control surface 152 as part of another wing lifting body 101. The thruster may be non-axially symmetrical in shape, and the control surface may be positioned exactly in the wake of said thruster 200. The thruster vigorously mixes a warmer motive stream provided by the gas generator, with the incoming cold ambient stream of air at a high drag velocity. Similarly, the mixture is sufficiently homogeneous to reduce the hot motive stream 600 from ejector temperature to a mixture temperature profile that will not affect the control surface mechanically or structurally. In this example, yaw may be controlled by changing the orientation of the control surface 152. The primary function of the thruster 200 is to generate thrust but also lift or attitude control. In this example, the yaw control is in direction 151 creating a rotation around the aircraft axis 10.

FIGS. 6A to 6C show the representation of FIG. 5 from different points of view.

For example, a pop-up jet having a rectangular pattern due to the rectangular exhaust plane of the propeller can also be vectored much more easily and in more directions than a propeller and an electric motor. In another example, a pop-up jet having a rectangular pattern due to the rectangular exhaust plane of the propeller is directed towards the leading edge of a short wing positioned some distance behind the propeller to maximize the lift benefit. The propeller can therefore generate the thrust necessary for the aircraft to move forward, in the direction primarily opposite to the direction of the jet outflow. Furthermore, the jet outflow moving at a faster speed than the aircraft speed and resulting from such a propeller or ejector will be used to increase the lift force resulting from its flow over the lifting body positioned behind such a propeller or ejector. The jet speed will always need to exceed the aircraft speed and the difference between the two speeds will need to be minimal to maximize propulsive efficiency. It follows that the greater the mass of flow providing thrust at lower speeds, even higher than the aircraft speed, the greater the propulsive efficiency. For example, using a propulsive efficiency equation known to those skilled in the art:

PE=2 V 0/(V+V 0)

where V is the propellant exit jet velocity and V0 is the aircraft airspeed, if the propellant jet velocity is 150% of the aircraft airspeed, the aircraft airspeed will be 50% of the propellant emerging jet velocity, and the propulsive efficiency will be 80%. After leaving the propellant exhaust section of an aircraft, the exhaust stream from most conventional jet aircraft is lost to the environment and the residual jet is not utilized, although the jet from, for example, a jet engine still carries energy in the wake. The exhaust flow is typically a round jet at higher velocities (and hence energy), mixing with a parallel flow at lower velocity and finally mixing with the aircraft exit vortex pair. Once it leaves the aircraft engine as exhaust, jet backflow no longer benefits the aircraft and the higher the velocity of the exhaust jet, the lower the propulsive efficiency and waste of energy to the environment.

One example utilizes the mixed stream emerging from the booster, which would otherwise be lost to the ambient in conventional aircraft, to generate lift or create directional change capabilities by directing it straight to a thin lifting body wing or other surface positioned directly behind said booster, for lift generation or attitude changes of the aircraft. Since the pressurized gas supply can be modulated or additionally utilized in a segmented manner through a network contained in the fuselage and wings of the aircraft, the drag and velocity of the efflux jet can be dialed in through primary or secondary methods. The primary method refers to pressure, flow, temperature modulation and/or segmentation (multiple supplies to multiple boosters distributed through the aircraft). The concept of segmentation involves the use of multiple propellant elements conveniently placed throughout an aircraft, i.e. segmenting the function of a single large propellant into multiple smaller ones which are being supplied with pressurized gas through a duct network. A secondary method may involve changing the geometry or position of the propellant relative to the neutral position of that propellant. For example, in level flight, supplying the appropriate gas pressure and flow to the propellant may result in jet outflow at 125% of the aircraft airspeed. In the case of a jet outflow axial velocity of 125% which is greater than the aircraft airspeed, the propulsive efficiency becomes 88%. If the emergent velocity becomes 110% at higher speeds with the same level of thrust generated by ambient air drag, then the propulsive efficiency improves to 95%.

Thrust and lift generator

Another example relates generally to a combination of thrust and lift obtained through a tandem system consisting of a thrust generating element directing a high velocity, non-circular outflow jet with velocity component in the primarily axial direction onto a thin lifting body located downstream of the outflow jet. The high local axial velocity of this outflow jet generates lift at levels considerably higher than the regular aircraft speed wing lift as ~ (Jet stream velocity)2. The outflow jet is a mixture of hot, high energy gases, provided to the thrust generator through ducts from an outlet of the high pressure gas generator, and entrained surrounding air. The entrained air is brought into a high kinetic energy level flow through a momentum transfer by the high pressure gases supplied to the thrust generator within the thrust generating element. The resulting air-gas mixture emerges outside the thrust generator and can be directed to point primarily in the axial, downstream direction toward a thin lifting body leading edge and/or the pressure side of the lifting body.

In most conventional aircraft, it is not currently possible to direct the jet outflow into a lifting body or wing to utilize its wasted energy. In the case of turbojets, the high temperature of the jet outflow effectively precludes its use for lift generation by means of a lifting body. Typical jet exhaust temperatures are 1000 degrees Celsius and sometimes higher when afterburning is used to increase thrust, as is true for most military aircraft. When turbofans are used, despite the use of high bypass in modern aircraft, there is still a significant residual element of non-axial steering, due to the rotation of the fan, despite the vanes directing the fan and core exhaust fluids mostly axially. The presence of hot core gases at very high temperatures and the residual rotational motion of the outgoing mixture, plus the cylindrical nature of the jets in the downwash, make the use of lifting bodies placed directly behind the turbofan engine impractical. Furthermore, the mixing length of the hot and cold streams from jet engines such as turbofans occurs in miles, not inches. On the other hand, the current use of larger turboengines generates large cylindrical downwash airflows the size of the propeller diameters, with a higher degree of rotational component velocities behind the propeller and moving large amounts of air at lower velocities. The rotational component makes it difficult to use the downstream kinetic energy for purposes other than propulsion, and therefore some of the kinetic energy is lost and not used efficiently. Some of the air moved by the large propellers is also directed into the engine core. In summary, jet outflow from current propulsion systems has residual energy and currently untapped potential.

In this example, the stream can be used as a lift generating stream by directing it straight towards a thin lifting body for lift generation. For example, when a jet outflow axial velocity is 125% greater than the aircraft airspeed, the portion of the wing receiving the jet outflow stream can generate over 50% higher lift for the same span compared to the case where the span is washed by the aircraft airspeed. Using this example, if the jet outflow velocity is increased to 150%, the lift becomes over 45% greater than the original wing at the aircraft airspeed, including a density droop effect if a pressurized blow-off from a turbine was used, for example.

Alternatively, a wing such as a light wing airfoil could be deployed directly behind the jet ejector exit plane, immediately after the vehicle has completed takeoff maneuvers and is transitioning to level flight, helping to generate more lift for less power from the engine.

Alternatively, using this example, the wing does not need to be as long in span, and for the same chord, the span can be reduced by over 40% to generate the same lift. In this lift equation familiar to those skilled in the art:

L = 14 pV2SCL

where S is the wing surface area, p is the density, V is the aircraft (wing) speed, and C<l> is the lift coefficient. A UAV with a wingspan of, say, 3.05 m (10 ft) can reduce the span to a mere 1.83 m (6 ft) provided the jet is oriented directly at the wing at all times during level flight, with a wing that is thin and has a chord, roll, and CL similar to the original wing. The detrimental impact of temperature on density is much less if the mixture ratio (or drag ratio) is large, and hence the jet is of only slightly higher temperature.

FIG. 7A depicts an alternative approach to having the jet engine placed on the wing and independently producing thrust. In FIG. 7A, the jet engine is no longer producing a jet outflow that pushes the aircraft forward, but is instead used as a gas generator and is producing a motive airflow to feed a series of ejectors that are integrated into the wing for forward propulsion. In this example, the gas generator (not shown) is integrated into the fuselage of the aircraft, and the green portion represents the inlet, gas generator, and ducts leading to the red ejectors, which are flat and, similar to flaps or ailerons, can be actuated to control the attitude of the aircraft in addition to providing the required thrust. FIG. 7A further depicts another (secondary) wing that is placed in tandem with the first (main) wing containing the thrust-augmenting ejectors, just behind said ejectors. The secondary wing is therefore given a much higher velocity than the aircraft's airspeed, and as such creates a high lift force since the latter is proportional to the airspeed squared. In this example, the secondary wing will see a moderately higher temperature due to the mixture of the propellant fluid produced by the gas generator (also called the primary fluid) and the secondary fluid, which is ambient air, entrained by the propellant fluid at a rate between 5-25 parts of secondary fluid for each part of primary fluid. As such, the temperature seen by the secondary wing is slightly higher than ambient, but significantly lower than the motive fluid, allowing the secondary wing materials to withstand and sustain lift loads, according to the formula: Tmix= (Tmotive+ER*Tam b) / (1+ER) where Tmix is the final temperature of the jet outflow fluid mixture emerging from the ejector, ER is the entrainment velocity of portions of ambient air entrained by portion of the motive air, Tmotive is the warmest temperature of the motive or primary fluid, and Tamb is the approaching ambient air temperature.

FIG. 7B depicts the front view of the aircraft shown in FIG. 7A with arrows illustrating the additional lift generated by the shorter tandem wings and the lack of engines on the wing.

FIG 7C depicts another example not falling within the scope of the claims featuring tandem wings. In this example, thrust-augmenting ejectors 701 that are part of the propulsion system are positioned on the main wings 703 (front wings) and are connected via ducts to and receive propellant fluid from a gas generator positioned within the fuselage. The ejector generates thrust and mechanically transmits the force to the aircraft. The efflux jet generates a high-speed steady flow that is used by the secondary wing (gray wing) 702 to produce additional lift. The combination of the two shorter wings produces more lift than a much larger span wing lacking the ejector thrust augmenters that rely on a jet engine attached to said larger wing to produce thrust.

FIGS. 8A and 8B illustrate yet another embodiment of the present invention. As shown in FIGS. 8A and 8B, the tandem thrust/lift generating system is attached to an air vehicle 804, where forward thrust augmentation ejectors 801, including leading edges and inlet portions for upstream air intake, are producing thrust just aft of a canard wing, with each of said ejectors positioned on the starboard and port side of the vehicle. The canard wing is oriented at a high angle of incidence and near stall when flying level, where the presence of the thrust augmentation ejector extends the stall margin of the canard wing 803. The thrust augmentation ejector 801 mechanically transmits the thrust force to the structure 804 and produces a downstream jet outflow consisting of well-mixed primary and secondary air streams, which in turn are used to generate significantly increased lift on the wing 802. The system is also replicated in the tail of the aircraft in a similar manner. The thrust augmentation ejectors 801 receive a compressor bleed stream from the gas generator 800, while the tail thrust augmentation ejectors receive the pressurized hot gases exiting the gas turbine of the gas generator 800. The combination of using compressor bleed air for the ejectors 801 and using hot exhaust gas for the tail ejectors as primary fluids, respectively, results in (1) increased thrust in horizontal flight due to ejector drag from ambient air and (2) additional lift generated on surfaces positioned aft of said ejectors, such as the wing 802 having leading edges. These elements positioned aft of the ejector are generally thin structures, and could be constructed from composite materials, including but not limited to ceramic matrix composites (CMC). This arrangement offers greater flexibility to change during the transition from takeoff to hover to horizontal flight and landing.

FIG. 9 provides additional detail to the tail (or hot) section of the illustration in FIGS. 8A and 8B. Thin structures 904 having leading edges are positioned in the wake of a set of hot thrust augmentation ejectors 901, which have leading edges and receive the primary (motive) fluid as hot exhaust gas from the gas generator 800, located near a cabin 805 and entrained air at the inlet 902 of the ejector 901. The duct connecting the exhaust from the gas generator 800 to the element 901 is embedded in the vertical fin structure 950. The ejectors 901 entrain incoming ambient air in the inlet areas 902 and eject a mixture of entrained air and motive gas at high velocity at the outlet 903 and primarily toward the thin tail structure 904, which in turn generates additional lift. Both elements 801 in FIGs. 8A and 8B and 901 in FIG. 9 may rotate about their main axis for VTOL and hover control. Additionally, each ejector of the ejector set 901 may rotate about the same axis with and/or independently of the other ejector.

FIGS. 10A through 10E show the various flows as well as lift versus angle of incidence with the point corresponding to the angle of incidence highlighted in each case. As the angle of incidence of a given lifting body increases, lift increases until boundary layer separation at the lifting body results in a stall just past the point of maximum lift (see FIG. 10D).

FIG. 10A demonstrates the lift and angle of incidence of the canard wing structure 803 illustrated in FIGS. 8A and 8B, at a zero degree (0°) angle of incidence, where the dot represents the lift force and the streamlines represent the flows around the raked face lifting body. FIGs. 10B through 10D show the result of the increasing lift force of the structure 803 as the angle of incidence or angle of attack increases up to the point of stall, depicted fully in FIG. 10D. Beyond the position of the lifting body (with respect to the angle of incidence) as shown in FIG. 10D, e.g., at the position depicted in FIG. 10E, lift rapidly decreases as the flow becomes turbulent, may separate, and the streamlines are no longer smooth. The lift increases almost linearly as the angle of incidence increases, but at the angle of incidence shown in FIG. 10D it reaches the maximum value, beyond which the flow separates on the upper side of the lifting body. In FIG. 10E, there is recirculation and increased drag, loss of lift generated by opposing flows, and onset of boundary layer separation. This causes the lift force to drop significantly and results in loss of lift.

FIGS. 11A and 11B show the lift characteristic curve of FIGS. 10A through 10E, with a second curve showing an extension of the stall range, demonstrating the improvement in lift versus angle of incidence beyond the stall point should the ejector be positioned relative to the wing so as to delay separation and facilitate ingestion of the boundary layer at high angles of attack. In FIGS. 11B, lift continues to increase without stall with angle of incidence, due to the presence of the ejector. Positioning the ejector beyond the apex of the lifting body allows for reunion or prevention of separation of upper boundary layer flow that would otherwise separate in the absence of the ejector ingesting said boundary layer, due to a high angle of incidence of said lifting body. The ejector is introducing a local area of low pressure at its inlet, forcing the ingestion of the boundary layer developed over the upper side of the wing lifting body. The stall margin becomes much greater by placing the thrust augmenter ejector beyond the apex of the canard wing or lifting body 803 of FIGS. 7C, 8A and 8B. These results indicate that the presence of the ejector extends the stall margin and allows greater lift forces to be generated by increasing the angle of attack beyond the value of the lifting body's stall angle of attack without the ejector being present. In addition, FIGS. 11A and 11B illustrate possible locations of the ejector with respect to the lifting body chord to re-streamline the flow around the lifting body.

FIGS. 12A to 12C represent another example that does not fall within the scope of the claims.

The main wing and thrust augmentation ejector system produce forward thrust and high velocity jet outflow conditioning that can be used for additional lift generation when coupled with a secondary lifting body (not shown, but could be placed in the wake or outflow downstream of the ejectors). As illustrated in FIG. 12A, the ejectors are formed by two (2) air curtain type halves, which together generate drag, momentum transfer and acceleration of ambient air using the primary or propellant fluid and expelling the final mixture of primary and secondary fluids at high velocities. The two halves 1201 and 1202 can independently rotate and translate into position on their own and relative to the wing such that they are optimizing the augmentation at any time, based on the aircraft attitude and mission (or point in mission), primary fluid condition (flow, pressure and temperature). This allows the throat formed by the two halves to have, in one case, a certain value, in yet another case, a larger or smaller value. For example, on takeoff, the two halves may point downward to allow the aircraft to take off vertically. The two halves may move independently and in position relative to each other to maximize thrust at maximum primary fluid flow rate and maximum drag flow rate, generating a certain inlet area to throat ratio that is favorable to maximize thrust. However, when flying level, the two halves of the ejector may instead be both horizontal and streamlined with the wing, with a smaller throat area for smaller pressures, temperature, and flow rate of the primary fluid, again maximizing thrust gain. The throat area, exit area, inlet area, and their ratios may also be tuned according to a thrust maximization algorithm. Both 1201 and 1202 contain a plenum chamber 1211 and 1212 respectively, connected to a conduit and receiving said primary fluid from, for example, a compressor bleed port of a gas generator. The two halves together form a variable inlet area 1201a and a variable outlet area 1201b and a diffusion shape formed by walls 1213 and 1214 respectively, to optimally spread the flow to maximize said thrust. The primary flow is introduced from plenum chambers 1211 and 1212 respectively into the throat area through multiple specially designed nozzles 1203 and 1204 respectively, in a continuous or pulsed manner.

FIG. 12C further describes the arrangement of this ejector for horizontal flight of an aircraft. FIG. 12C shows that the flat ejector can be inserted into the thickness of the wing's lifting body when all elements described in this disclosure are used for highest efficiency. FIG. 12C shows the outline of said inner and outer ejector surfaces and FIG. 12B shows the 3D model of the lower and upper halves 1201 and 1202 of the described flat Coanda ejector, integrated with the wing. The two halves, which can be actuated independently, together form an inlet 1201a and an outlet 1201b; they allow the high-speed introduction of a primary fluid through the primary nozzles 1203 onto the Coanda surfaces 1204.

FIGS. 13A through 13C depict how the present invention can control pitch, roll, and yaw of the aircraft using the thrust augmenting ejectors in conjunction with thin lifting bodies positioned in the wake of the ejectors. With respect to pitch, the hot and cold ejectors can be independently rotated about their major axis to cause the aircraft to pitch forward or backward. Pitch control is affected by ejector thrust splitting forward/aft and/or modulation of the flow of the propellant fluid supplied to the ejectors. With respect to roll, the ejectors can be independently rotated to cause the aircraft to roll. With respect to yaw, a combination of additional rotation about a perpendicular axis with positioning of the thin lifting bodies in the wake of the jet outflow can be used to cause a change in aircraft attitude. This embodiment of the present invention allows these maneuvers with the use of special joints that can tilt, transmit loads and allow the passage of primary fluid to said ejectors.

Coanda Device

In yet another example not within the scope of the claims, the propeller and/or thrust generator of the tandem system have the ability to entrain large amounts of air and accelerate it to jet outflow velocity. This is achieved by employing a Coanda device.

These flow augmentation devices have been generally described by different publications which will be discussed in greater detail below. For example, in his paper "Theoretical Observations on Ust Augmentation", (Reissner Anniversar y Volume, Contributions to Applied Mechanics, 1949, pp. 461-468), von Karman describes in great detail why a Coanda device results in significantly greater thrust augmentation across multiple jets. Similarly, U.S. Patent No. 3,795,367 (Mocarski) discloses a device for air entrainment with high augmentation ratios exceeding 1.8, while U.S. Patent No. 4,448,354 (Reznick) applies a linear Coanda device to the VTOL capability of a jet engine. In these aforementioned publications and other references not mentioned herein, the application of Coanda devices has been limited and described only for VTOL and not for horizontal flight. A main learning was that scalability and application for horizontal flight was not practical, particularly for Coanda-type axisymmetric devices where their size would induce an increase in aerodynamic drag for larger aircraft. However, an application for a small UAV may be more suitable with a higher degree of integration. The examples of the present disclosure are capable of integrating the ejectors with the fuselage and engine or propulsion system because the vehicle does not need to consider a large seating capacity. Integration as described in these examples is not currently practical or commercially reasonable in large commercial flights.

This example improves the Coanda device and applies it using novel techniques for improved entrainment and retardation or separation avoidance at its aggressive turns within the device. Although the compactness of these devices is critical for their deployment in aviation and other fields, the inlet portion needs to be large to improve air entrainment. Reznick argues that a circular element is more efficient than a linear one. Mocarski shows that entrainment is critical for thrust enhancement. The diffuser portion must be long enough to ensure that boundary layer separation does not occur within the device and mixing is complete at the exit of the device. Conventionally, these diffusers have been long with a very gentle slope to minimize the risks of boundary separation. The present disclosure shows improved entrainment in the devices by means of novel elements that are based on 3D geometric and fluid flow effects and the utilization of separation avoidance techniques in the Coanda device. The preferred example has a entrainment ratio between 3-15, preferably higher. In another example, the device will receive the propellant gas from a pressurized source such as a gas generator, a piston engine (for pulsed operations), or a compressor or supercharger. Another feature of the present disclosure is the ability to change the shape of the diffuser walls of the flat ejector used for propulsion by retracting and extending the surfaces to modify the geometry so that maximum performance is obtained at all points of the aircraft mission. In addition, the need to rotate the entire ejector 90 degrees is no longer necessary for VTOL and hovering, when the fully deployed diffuser walls are used to direct the jet efflux downward.

Another example not within the scope of the claims introduces fin-like elements into the diffuser walls of a Coanda ejector which is divided into 2 halves, illustrated in FIG. 14, as upper 1401 and lower 201 ejector means which are each similar to an air curtain. Elements 115 and 215 are actuators or linkages, which allow movement of said surfaces to the desired position of the diffuser at both 110a and 210a, respectively.

Yet another example not within the scope of the claims discloses how staggered 3D inlet geometries and/or 3D primary fluid slot features, either independently or working together, significantly improve propellant performance, along with the introduction of flow separation avoidance patterns into the propellant. For example, as illustrated in FIGS. 15A through 15C, the 2D inlet is replaced by a 3D inlet. FIGS. 15A through 15C further illustrate the multiple 3D elements of the disclosed ejector, improving its performance over the baseline, and 2D ejectors having the inlet, throat, and diffuser in the same planes, respectively.

The inlet may further match the boundary layer profile shape formed behind the apex of an aircraft main wing lifting body (as illustrated in FIG. 16A), thereby helping to ingest the boundary layer and delay overall stall (improving across all margins), further illustrated in FIG. 25 for the position relative to the lifting body. FIGs. 11A and 11B illustrate the benefits of positioning it as such, relative to the lifting body and its boundary layer profile.

FIG. 16A shows the flat ejector and a wing structure to improve its high angle of incidence performance and stall margin. The ejector is supplied with a primary fluid from, for example, a gas generator, and is positioned to make the flow over said lifting body more efficient in delaying stall.

FIGS. 16B through 16D show different angles of the illustration shown in FIG. 16A, with details of the positioning of the ejector on the wing, the plenums supplying primary fluid to the ejector, and their relative position to each other and the lifting body.

The ejector depicted in FIG. 14 is of planar geometry and contains an upper and lower portion, both of which introduce the motive fluid as wall jets into a multitude of slots and generally perpendicular to the jet outflow direction of the flow or streamlines, elements 1401 and 201 which can rotate independently about axes 102 and 202. The curved walls called Coanda walls 104 and 204 allow the primary jets to follow the curvature and entrain in the process and in a ratio greater than 3:1, secondary air, generally arriving from the flow above an airfoil such as a wing upper surface boundary layer. The primary nozzles 103 and 203 are of various shapes with various 3D effects to maximize the drag ratio, such as mini delta wings 212 in FIG. 22B, or they can be fluidic oscillators fed by said plenums fed with motive fluid to generate a pulsed operation of jetting motive fluid onto the Coanda walls. The mixed fluid reaches the throat area (minimum ejector area) in a pure axial direction. Beyond this point, the present invention introduces a segmented movable diffusing section, such as a fin, only that it has a main role in the performance of said ejector by vectoring and/or maximizing its performance.

For example, at takeoff, the inlet of said ejector is fixed and still above a forward-pointing lifting body 1700 in FIG. 17A. FIG. 17A depicts the deployment of said ejector formed by an upper half-ejector (1401) and a lower ejector (201) and in conjunction with the main wing of a drone aircraft. The two half-ejectors can rotate around the axes 102 and 202 respectively and can also translate according to mission requirements. FIGs. 17B and 17C show the case where only the upper half-ejector 201 is actively used with a primary fluid, while 1401 is replaced by a simple flap. As before, 201 can rotate around the axis 202 and translate with respect to the axial position. Semi-ejectors receive the primary fluid under pressure from, for example, a gas generator such as a gas turbine and allow it to pass through primary nozzles, which may employ fluidic oscillators (i.e. pulsating at certain frequencies such as up to and including 2000 Hz to generate pulsating entrainment of the secondary flow).

In FIG. 14, the upper ejector diffuser means 210 extends to form a curved surface 210a and guide the mixed primary and secondary flows downward. At the same time, the lower diffuser 110 also extends at 110a, maintaining the appropriate ratio of area growth and mixing characteristics to obtain the maximum thrust required by the aircraft. Portions of 110a and 210a may not deploy and 110 and 210 are still independently controlled according to an appropriate program. In addition, the upper element 201 may or may not axially move to follow mission needs. In one example, different amounts of primary fluid and/or delivered under different conditions may be delivered to the upper 201 or lower 1401 elements in a continuous or pulsed manner. The diffuser surfaces 110a and 210a may contain dimples and other elements that retard or prevent boundary layer separation. Additionally, the secondary nozzles may also be opened if the fully extended 110a is used, at particular and potentially staggered locations and may be pulsed according to the fluidic oscillator operating modes to provide a pulsed mode of operation to the ejector.

When fluid is received from a compressor bleed, the motive air is at a lower temperature. The exhaust gas from the hot end of the gas generator (turbine exhaust), for example, for motive gas temperatures of 815 °C (1500 °F) at a pressure of 30 psi compressor air discharge and drag ratio of 5:1 and ambient temperature of 38 °C (100 F), the mixture temperature becomes 180 °C (335 °F), for which the air density is 1.6E-3 slugs/ft3 or 0.84 kg/m3, a -30% drop from ambient. As such, the overall wing span can be reduced by -10%, even accounting for density reduction effects, when a lifting body is deployed behind the main booster. For drag ratios of 10:1 (better than the 5:1 design), for similar conditions and a jet washout of 125% of the aircraft airspeed, the lift benefit is greater because the mixture density is now higher, at a mixture temperature of ~200F, and the lift generated over the jet-washed span is -16%. In this example, the wing can be reduced in length accordingly.

Thrust generator

Another example not within the scope of the claims relates generally to a novel 3D thrust generator that is capable of receiving pressurized gases from a plenum, entraining ambient air under stationary or moving conditions (including but not limited to those conditions greater than 0.05 Mach), accelerating the air through momentum and energy transfer with the high pressure gases, and directing the well mixed fluids into a high velocity, non-circular efflux jet with a primarily axially directed velocity component. The efflux jet may be a mixture of hot, high energy gases, provided to the thrust generator through ducts from a high pressure gas generator outlet, and entrained air at ambient temperature. The entrained air may be brought into a high kinetic energy flow through momentum transfer with the high pressure gases supplied to a propulsion device, within the thrust generator. The resulting air-gas mixture exits the thrust generator and points primarily in the axial, downstream direction opposite the direction of vehicle trajectory. The well-mixed stream provides a mostly unidirectional stream of cooler gas at high velocity, which can be used for propulsion, hovering, lift generation, and attitude control via lifting bodies placed in the cooler jet wake. This is not seen on any conventional jet fuel engine-powered vehicle. This thrust generator may be independent of the fuselage, embedded with the fuselage at the front or rear of the vehicle, and/or embedded in the wings for stall margin enhancement.

Reznick invented a circular device with the primary nozzles being separated from the Coanda surface and thus not generating wall jets. While Reznick teaches that additional secondary fluid is being admitted due to displacement to the Coanda surface, his application, however, is strictly circular in shape and thus cannot be scaled into a more practical application for aircraft of larger flows and, for example, still integrated with a wing, as the aerodynamic drag becomes increasingly larger. Furthermore, the slots also appear to be simple in geometry and do not present any particular 3D features for mixing enhancements. The present disclosure introduces an aerodynamic thruster that generates a rectangular-shaped outflow at the trailing plane, to use the energy for additional lift generation on the thin lifting body, an improvement and departure from the circular application of Reznick, which cannot be effectively used along a longer lifting body for lift generation in flight level other than its own diameter, and cannot be deployed over a wing to ingest the boundary layer of a wing.

Primary nozzle geometry

It is noted that in all the patents described the inventors are not employing any feature that increases the area of the primary jet to the secondary flow, and therefore there are limitations of the inventions described. Furthermore, there is no staggering of primary nozzles in the Coanda device, except for the Throndson presence of the central primary nozzles, which are not placed in the Coanda device but in the centre of the inlet perimeter of the Coanda primary nozzles. The primary nozzles are therefore generally placed in the same axial plane and not staggered, nor are they different in size from the adjacent ones, but of the same size and shape. If for a circular Coanda device this is optionally advantageous, for a non-circular one having a constant spacing between opposite sides of the primary Coanda nozzles along the length of its largest dimension of the inlet plane, the thrust resulting therefrom would be equally distributed in an ideal situation but during horizontal flight, if such a device is employed for thrust generation, the secondary incoming air would be unevenly admitted into the device and therefore thrust generation will impose challenges to the wing structure and its design. This is mainly because in the above-mentioned prior art, these devices were envisioned to be used in the initial and final stages of an aircraft's flight and not as a sole thrust generating propellant for the entire mission, from take-off to landing and including hovering and horizontal flight. Indeed, Throndson's invention is applicable only for vertical take-off and landing and hovering, the powerplant taking the horizontal flight function of providing thrust via a turbojet or turbofan. Thus, in his invention, devices including the Coanda ejector are turned off and form the wing's lifting body in horizontal flight, i.e. not operational or active during horizontal flight, after the transition from takeoff. On the other hand, Reznick teaches a circular device with primary nozzles for thrust augmentation but without embedding it with a wing for horizontal flight and exploiting both the inlet and outlet of the device so that it does not generate thrust.

FIGS. 18A through 18D show conventional arrangements for Coanda-type ejectors. FIG. 18A depicts a traditional prior art circular Coanda ejector. FIG. 18B shows a prior art flat wing-embedded Coanda-type ejector. The source of the primary fluid is a gas turbine and the ejector is advantageously intended, eventually, to be used for vertical takeoff and cutting in horizontal flight. FIG. 18B encompasses the elements disclosed as Throndson variables, including diameters, angles and lengths.

FIG. 18C is Reznick's Figure 3 and shows another circular example where hypermix nozzles are employed and the primary fluid nozzles are set back from the ejector walls. The primary jets are therefore no longer wall jets. Reznick only covers circular ejector geometries, clearly intended to be used for takeoff assist due to scalability limitations.

FIG 18D depicts an example not falling within the scope of the claims with circular Coanda nozzle elements, with a plenum chamber 211 supplied with primary fluid, which is accelerated through the primary nozzles 203 and injected as wall jets onto the surface 204.

Throndson utilizes a non-circular ejector shape but also rectangular slots. A rectangular slot is useful in such an application but produces limited surface area for oncoming secondary air shear jet entrainment. In fact, a rectangular slot described by the inventors above produces a jet entrainment characteristic for a rectangular slot perimeter of the given dimensions, 2L+2h=2(L+h) where L is the length and h is the height of each slot. A much greater amount of secondary flow is entrained if a larger primary nozzle perimeter is used, including the impact of 3D features. Axially staggering the vertices of a zigzag or wavy (sinusoidal) wall of a primary nozzle as shown in FIGS. 18A through 18D greatly improves secondary air entrainment as taught in the present disclosure. A pulsed operation by embedding fluidic oscillators with the primary nozzles further improves the efficiency and drag characteristics of the ejector.

FIGS. 19A through 19D depict some of the proposed changes to the primary nozzles for improved performance. FIG. 19A shows a zigzag configuration of primary nozzles along the circumference of the ejector inlet area, while the perimeter of the primary jet exposed to the secondary flow is doubled compared to a single slot perimeter, thereby increasing drag through turbulent shear layers developed between said zigzag walls of the primary nozzles. FIG. 19B shows a rectangular slot with increased roughened perimeter to generate additional turbulence, and thus increase drag 1.5-4 times compared to the original smooth walls of a rectangular slot. FIG. 19B schematically depicts the increase in the area of the primary jet surface to secondary or entrained air, with the 3D structure of the barbs explained in axial direction. While normally in a rectangular slot arrangement the secondary air is entrained primarily between two adjacent slots and on the side of the outer radius slot, entrainment is now greatly enhanced through surface and 3D effects. FIG. 19C explains the primary and secondary jets respectively, with the turbulence generated by the 3D features greatly enhancing the mixing and momentum imparting from the primary flow to the secondary flow over a shorter distance. FIG. 19C shows the interaction and resulting flows of such adjacent rough walled slots, with the red arrows representing the primary fluid and the blue arrows the entrained secondary fluid. Shear layers form along the walls and the increased perimeter results in significantly greater entrainment of secondary flows for the same inlet primary flow conditions. Pulsed operation of the primary nozzles further improves the entrainment ratio.

Another feature of the present disclosure is the introduction of advantageous features within the primary nozzle (see FIGS. 19E and 19E). It is well known that a flow over a delta wing produces vortices that are opposite in direction toward the center of the delta wing. A miniature feature is placed on some or all of the primary nozzles to generate such vortices emerging from the primary nozzle. In this case, the vortices advantageously entrain significantly greater amounts of secondary air into the ejector, improving its mixing and the transmission of momentum carried by the primary fluid exiting the primary nozzles in a continuous or pulsed manner.

FIG. 19E depicts the flow over a delta wing obstruction positioned within the primary nozzles at its center, which changes the flow pattern such that it significantly increases the drag ratio of a normal primary nozzle slot, without requiring changes in pressure flow rates and temperatures. In particular, the primary vortex cores are opposed in a direction toward the center of such a slot or delta wing, entraining significant secondary fluid from the area between the slots.

FIG. 20 depicts the thermodynamic cycle of the present disclosure with the evolution of the working and entrained fluids to obtain high thermodynamic efficiency. FIG. 20 demonstrates that air entrainment will determine the movement of point D representing the mixing condition between primary gas and secondary air, in a propulsion thermodynamic cycle diagram, to the left and to lower values of temperature and entropy. This is optionally advantageous for a high propulsive efficiency device, where massive amounts of air are entrained and accelerated to relatively lower exit jet velocities, while maintaining a high thrust level due to higher mass flow rates, a key ingredient to realize high propulsive efficiency. In FIG. 19A, an increase in perimeter by a factor of 2 is shown by replacing each length of a normal rectangular slot perimeter by an equilateral triangle of 2x longer perimeter in the same plane. The perimeter may be further increased by staggering all of the slot wall apices in various planes (see FIG. 19B). The result of such a primary nozzle is to increase the amount of fluid entrained as secondary fluid by at least 15-50% by intimate mixing within the formed shear layers. If the initial secondary air condition was low velocity, then the performance of the rectangular and non-rectangular perimeter shapes may not be much different, however, when the ejector is moving forward and the oncoming secondary air velocity is significantly higher, such as between Mach 0.0 and Mach 0.25, then the spiked profile shape of the primary nozzle may also be significantly improved by placing the innermost and outermost spikes of the primary nozzle forward and aft of the axial plane of said rectangular slot. In other words, each primary nozzle now becomes a 3D structure that will delay or anticipate the entrainment of secondary air in an efficient manner, thus improving the overall entrainment rate. In a Coanda ejector it is optionally advantageous for the entrainment of secondary air and mixing with the primary air for momentum transfer to occur quickly and over a short distance. Adding this and other 3D elements to the primary nozzle helps improve the performance of such an ejector.

Another feature related to the primary nozzles as employed in this example is the introduction of fluidic oscillators within the flow path of the primary nozzles.

These fluidic oscillators provide, for example, switching of up to 2000 Hz between two adjacent primary nozzles to alternate wall jet flows and improve entrainment rates through pulsed operation of the motive fluid.

Yet another feature implemented in this invention is the staggering of the nozzles with their features, by being placed at various locations along the Coanda surface and thus introducing the primary flow at multiple axial locations, adjacent to the wall in a wall jet manner and in a pattern that increases entrainment and mixing of the secondary fluid. For example, FIG. 21 shows such an example where a V-shaped vortex generating feature is staggered compared to a normal rectangular slot and injecting at least 25% of the total primary fluid before the balance of the primary fluid mass flow is injected at a later time. This injection before the rectangular slots results in a higher entrainment rate sufficient to increase ejector performance significantly. Furthermore, in FIG. 21, the nozzles 205 inject the primary fluid before the primary nozzle 203. The nozzle 205 has a feature that introduces a more favorable entrainment of the secondary flow through shear layers and these nozzles 205 are staggered both axially and circumferentially compared to the primary nozzles 203. The primary nozzles 203 have a delta wing feature that is provided with a support leg connected to the midpoint of the primary slot structure at the innermost side thereof and having the delta wing structure pointing against the primary fluid flow to generate 2 opposite vortices in direction and strongly entraining from both sides of the primary nozzle 203 the already entrained mixture of primary and secondary fluid flows resulting from the nozzles 205. The vortices and the V-structure of the primary nozzles result in a 10-100% entrainment improvement compared to the primary nozzles 203. non-staggered rectangular slots and an overall improvement in momentum transfer from primary to secondary flow.

Furthermore, FIG. 21 depicts a simpler construction using delta fins arranged within smooth wall slots to form specific delta wing flows and shear layers which advantageously increase the drag ratio more than 2 times compared to smooth wall rectangular primary slots. All of these elements can be combined for the best drag ratio. The present disclosure improves the surface for flow separation retardation through elements 221 by placing dimples on the Coanda surface 204, where said Coanda surface 204 has a relatively aggressive twist for the shortest distance change from the primary flow direction, originating radially from the primary nozzles 203, to the axial direction opposite to the thrust direction, towards the throat 225. The dimples prevent flow separation and significantly improve the performance of the ejector, along with the delta turbulators shown in FIGS. 19D and 19E.

FIG. 23 illustrates certain features according to an example not falling within the scope of the claims. In particular, FIG. 23 compares a similar primary slot height used by Throndson to the ratios used by Throndson to demonstrate the improvement of this example. The radius of gyration at the slot height of this example is below 5:1 with improved separation retarding dimples placed on the Coanda surface. In FIG. 23, the radius R' is approximately 2-3 times smaller than the radius R of the Throndson patent, for similar slot heights. It follows that a ratio less than 5:1 is possible, due to the use of a logarithmic profile along with the employment of dimples on the curved Coanda surface, to more aggressively return the flow from purely radially at the exit of the primary slot to purely axially at the throat. As a result, much faster turning occurs without flow separation, such that the throat of the device can be made larger than that specified by Throndson by at least 25-100%. The half angle of the diffuser can also be made significantly more aggressive than in the prior art, allowing for a much shorter diffuser to be implemented and faster momentum transfer between the primary and secondary flows. As such, FIG. 23 highlights the differences between the present disclosure and the prior art, especially where more aggressive elements such as turbulators, primary nozzles, dimples and moving walls improve upon the prior art.

Coanda ejector

In general, the design of a Coanda ejector applied to an aircraft has been described by many publications. For example, U.S. Patent No. 3,664,611 (Harris) teaches a Coanda-type ejector embedded in the wing for vertical takeoff and landing purposes. The device is not active during cruise, see FIG. 24. Harris is silent on the use of efflux to generate additional lift in a tandem-type arrangement. Furthermore, Harris does not apply the device for use in horizontal flight conditions. Instead, according to conventional practices, the device is collapsed on a lifting body wing during horizontal flight conditions.

Mocarski, on the other hand, teaches that in a Coanda ejector, the high energy primary fluid, also called the motive fluid, is injected as a wall jet and the principle of such a device is to determine a low pressure zone where ambient air is entrained, followed by a mixing and convergence zone towards a throat, followed by a diffuser to expand the mixture back to ambient pressure at high velocity. U.S. Patent No. 3,819,134 (Throndson) modifies and improves this concept described in Mocarski.

Throndson describes an improvement of the technology by adding a primary flow to the centre of the Coanda-type ejectors to further entrain secondary fluid and to improve nozzle performance, with the central primary nozzle using 30-70% of the total primary fluid and the remainder being used in the parametric Coanda-type nozzles. Throndson claims that the thrust increase is greatly improved by this combination, and does not comment on the geometry of the primary fluid nozzles, which appear as simple slots or holes. Furthermore, the slots appear to be continuous or discontinuous and featureless. Throndson says nothing about using the efflux jet for downstream lift generation, and in fact only employs the device for take-off and transition and landing, not for cruise conditions, much like Harris.

The present disclosure further improves the Coanda ejector by generating thrust in all flight conditions by rotating the Coanda device and placing a thin lifting body in the jet stream thereby generating more lift. Such turbofans generally employ at least two advantages over Harris and Throndson:

First, the use of the ejector downstream of a wing such that it ingests its detrimental boundary layer under cruise conditions, improving the wing's aerodynamic performance and allowing high angles of incidence, thereby increasing its overall performance. The present disclosure also allows operation of the ejector in all phases of flight, from takeoff through hover, transition, cruise and landing. One example also allows the use of a half-ejector (1/2 of a flat ejector as described by Throndson or Harris) in conjunction with a wing slat to form an asymmetric Coanda ejector with boundary layer drag only at the outer edge of said boundary layer and forming a diffuser with said wing slat, including vectoring the pusher by moving the slat and the air curtain-type Coanda ejector in coordination for takeoff and landing.

Secondly, by using a thin lifting body downstream of (at the ejector wash) in minimum horizontal flight but also for other flight conditions for additional lift generation in higher velocity flow, it allows the lifting body and propulsive tandem to be more compact and efficient while generating considerable lift, compared to those disclosed in the above mentioned patents. In this example which does not fall within the scope of the claims, the shape and profile of the propulsive efflux jet is critical to achieving its novel efficiency and functionality. Said thin lifting body is placed at a convenient distance from the exit plane of said ejector/propulsor in order to maximize lift but also before the efflux jet energy is dissipated to the environment. This is convenient and practical since the energy of any jet propulsion device is normally dissipated only over very long distances behind the aircraft.

It is also important to understand that both elements of such a tandem need to work together in an efficient and optimized manner, including movement at certain angles and speeds favorable to the concept. The pusher/propeller mechanically transmits the thrust component to the aircraft fuselage or its main wing, while the thin lifting body downstream of the propeller is in mechanical contact with the fuselage and not the propeller, still receiving its efflux jet in such a way as to maximize the aircraft's lift and allow maneuvering through movements of certain surfaces on said thin lifting body.

Another feature of the present disclosure provides the ability to use the same nozzles to lift the aircraft in hover and landing as well as for cruise purposes. The lift system described in U.S. Patent No. 8,910,464 (Ambrose) represents the common mainline of VTOL jet fighter aircraft. It has limitations due to the additional weight carried in cruise mode, namely the lift fan and its auxiliaries. Under current VTOL technology, the cold nozzles (forward nozzles) and lift fan are turned off during horizontal flight leaving the main exhaust nozzle to provide the reaction force to propel the aircraft forward under forward motion conditions such as cruise. An example not falling within the scope of the claims combines Coanda nozzle thrust generating elements with the aircraft propulsion system allowing the ejector to be employed in all stages of flight with weight minimization and elimination of moving parts. It also allows the use of such an ejector to minimize aerodynamic drag and maximize lift in a unique way during horizontal flight.

Mocarski presents the same technology for a Coanda device with continuous or discontinuous primary fluid slots, mainly circular or linear. In all these patents, the Coanda surface is a circular or smooth 2D profile to determine a simple bonding of the boundary layer without particular elements that can enhance drag, increase the aggressive twist of the Coanda surface or delay its separation. In Coanda type ejectors it is critical that the surface twist allows the boundary layer of the wall jet to grow and mix with the secondary air and not separate. Once the primary flow jet from the primary nozzles separates, the Coanda ejector will not operate efficiently or at all. Therefore, it is paramount that the surface curvature is such that it allows maximum growth of the boundary layer and entrainment of the secondary fluid and its mixing with it, without separation at the wall.

On the other hand, if the curvature is too large, the device becomes impractically long and large in diameter, also restricting the amount of secondary fluid drag and mixing and inducing a very long diffusion part of the device. The ratio of slot to radius of the Coanda turn is described by Throndson to be between 1:5-1:15, but a ratio less than 1:5 should be ideal for a fast turn. The Coanda turn is clearly stated by Throndson to be ideally between 30 and 110 degrees compared to the axis of the device. If the diffuser section is being made too large, this is a major limitation in deploying the technology for an aircraft in horizontal flight, as the length of the diffuser would impose significant additional drag and weight on the aircraft. If the twist becomes >110 degrees, then the diffuser can become shorter and enhance mixing over a much shorter distance, ensuring intimate mixing and transfer of energy and momentum to the secondary flow prior to the mixture exiting the device. It is noted that the diffuser walls are also planar and free of 3D elements to enhance the mixing process. An example not falling within the scope of the claims introduces movable walls beyond the throat section, especially in the diffusion area of the ejector, in a manner favorable to vertical takeoff and landing of an aircraft, without the need to move the entire ejector around its horizontal axis but rather by extending the segmented diffusing surfaces in a manner described below.

Coanda surface

The Coanda surface, as taught by Reznick, Mocarski and Throndson, should be a round curvature, with Throndson providing even finer details than the slot height to radius ratio range of 1:5 to 1:15. A logarithmic profile is preferred by those skilled in the art as it provides the most rapid boundary layer growth without jet wall separation. However, one example not falling within the scope of the claims achieves much more aggressive twisting by introducing dimples into the Coanda surface to significantly enhance surface twisting in order to hold the flow together while mixing and moving the mixture into the throat and diffuser. Aggressive twisting is preferable because it allows the ability to rapidly mix and twist the flow in the axial direction, through the throat and into the diffuser section. The twisting of fast moving fluids, in fact, can hold the boundary layer together while the boundary layer grows and mixes with the central flow.

The dimples may be of different sizes, may be staggered or aligned, may be located in areas where the fluid spin is more aggressive and not in areas where the fluid spin is less aggressive. The dimples may also be employed in a more aggressive diffuser, where the diffuser half angle is not constant but variable, increasing and then decreasing to 0 as represented by element 105 in FIG. 14.

FIG. 14 depicts one of the improvements achieved when compared to Throndson. FIG. 14 compares a similar primary slot height used by Throndson with the relationships provided in Throndson. In particular, FIG. 14 shows the upper and lower halves of the ejector halves which together form a better, more flexible and efficient ejector that can be suitable for both vertical takeoff and hovering flight and for horizontal flight under cruise conditions. The lower wall (1401) of the ejector halves more aggressively utilizes the primary nozzle wall jets to more aggressively spin the flow around axis 102 and onto surface 103, a Coanda inlet surface. The maximum height point of this curve is axially positioned at a point at a distance of approximately G from the similar lowest position of the curved wall (closes to the axis shown in blue) of element 201. Therefore, the two half-ejectors (or air curtain ejector walls) 1401 and 201 are staggered, i.e. their inlets are not axially positioned at the same location. Similarly, the axial location of minimum distance of 1401 and 201 being staggered by the distance 'G', their diffusers 110 and 210 can be changed in shape by means of actuators 115 and 215 respectively, where the flat segmented surfaces forming 110 and 210 are converted into a curved cross section 110a and 210a respectively, directing the flow within said ejector downward or in various directions as the mission dictates. FIG. 14 also represents the changes in relationships compared to the prior art.

Furthermore, the gyration ratio at slot height of the present invention is below 5:1 with enhanced separation retarding dimples placed on the Coanda surface. As a result, much faster gyration occurs without flow separation, so that the throat of the device can be larger than that specified by Throndson by at least 25-100%. Furthermore, by applying a constant variation of the half angle of the diffuser portion (i.e. non-linear growth of the wall away from the centerline) and employing dimpled surface on said diffuser, its dimension can grow without flow separation more aggressively, resulting in a shortening of the overall length of the device.

Furthermore, if both the upper and lower halves of the ejector act separately with respect to fluid delivery and functionality, but can work together to entrain, mix and spread the mixture to the exit plenum, then performance is greatly enhanced by the additional diffuser movable walls on both the upper and lower surfaces of the flat diffuser. This in turn also allows the more compact device to be implemented in conjunction with a wing structure for propulsion reasons in horizontal flight or vertical takeoff, hover and landing without the need to rotate the entire structure.

Furthermore, the use of dimples allows for a variation of the wall from the initial point to the exit around the entire perimeter of said ejector, thus allowing for good integration with the wing structure. A different structure is used at the round ends of the ejector, where dimples or special features on the primary nozzles may not be required for the ejector to function satisfactorily. FIG. 22A shows an ejector having significant 3D features as described in this invention. Furthermore, in the more aggressive areas of the turns of the lower wall of the diffuser, the use of dimples (item 221 in FIG. 21) will also allow for an upward turn up to 90 degrees downward or even more. This is more aggressive than the prior art (e.g., Throndson; see Fernholz, H.H. "Z. Flugwiss. 15, 1967, Heft 4, pp. 136-142). FIG. 14 shows a cross section of an ejector having significant 3D features as described in this disclosure. FIG. 14 also shows mainly segmented walls in the diffuser, capable of redirecting the ejector jet and maximizing its performance from changes in the diffuser areas and mixing zones. The ejector inlet plane is not located in a plane (non-planar), and therefore it is possible to place the ejector above a wing structure so that the ingestion of the boundary layer improves the performance of the wing's lifting body, as can be seen in FIG. 14 with the dimension G (clearance) between the two inlets of the upper and lower halves of the ejector. The Coanda surface (103 in FIG. 14; 204 in FIG. 22A) therefore does not support the primary fluid wall jets in the same axial position, but rather earlier in the vicinity of the airfoil surface and later away from the airfoil surface of the wing, in the axial time history direction.

FIGS. 22A through 22F illustrate different examples not falling within the scope of the claims, where 3D features are used at the inlet. FIG. 22F presents an upper ejector half.

FIG. 22B shows element 212 as a delta turbulator placed within a primary nozzle, which greatly improves entrainment of secondary flows, and dimples such as 222 which improve coupling and prevent separation even at the most aggressive turns of the Coanda curved wall 803 of FIGS. 22A through 22F, and FIG. 14.

In FIGS. 22A-22F, element 212 is also introduced into the primary nozzles to improve drag and may or may not be employed on the opposite side of the ejector depending on conditions and to improve performance. Dimples 221 are placed in the contour 204 and diffuser to ensure good momentum transfer over the shortest length and to generate as uniform an exit velocity and temperature profile of the primary and secondary fluid mixture as possible and prevent flow separation. These primary fluids may be pressurized air from a compressor bleed or pressurized exhaust gas from a gas turbine or a mixture of both and may be fed to the upper and lower halves of ejector 1401 and 201 separately, adding another degree of freedom in maximizing thrust generation efficiency.

In FIG. 22A, 3D the features increase the perimeter exposed to the flow and allow for a higher entrainment ratio. In FIGS. 22B and 22C, specially designed turbulators, such as a delta wing positioned at the center of the primary slots, cause the flow from the primary fluid plenum chamber, constantly supplied by, for example, a gas generator, to be accelerated in a passage and forced to flow over said delta turbulator 212. The element 212 forces the flow into patterns that greatly enhance secondary flow entrainment through a number of mechanisms including shear layers, rotating and counter-rotating turbulent flows, and increased wetted perimeter of said primary nozzles 203. Embedding fluid oscillators within the primary nozzles also provides additional capacity for entrainment through pulsed operation in adjacent primary nozzles.

FIG. 25 shows an ejector arrangement with the advantage of having a 3D inlet with the lower lip (22) of the ejector being closer to the upper airfoil surface wall side 20 and beyond the apex of said airfoil surface, axially stepped and positioned ahead of the upper lip (23) of said ejector further from the airfoil surface 20. The position of the lips is modeled to match the most probable boundary layer velocity profile (21) resulting from the airflow near the airfoil.

By anticipating the drag of the airflow closer to the airfoil wall with the lip 22, as compared to the drag of the boundary layer lip 23, a better distribution in inlet and ejector performance is obtained. In one example, the ejector can be moved up and down (in the vertical direction to the upper airfoil wall) to optimize performance. This will allow for better airfoil performance at higher angles of attack and better performance of the ejector itself, through improved drag processing. 3D elements that differentiate this disclosure from the prior art include the position of the inlet lips 22 and 23, the relative position of the curved walls 204, the positioning of the throat area 24 and the diffuser walls 25. In one example, the two halves of the ejector can be moved independently relative to each other and the airfoil, resulting in a consistently optimized position relative to aircraft performance.

FIGS. 17A to 17C show a flat ejector placed on the lifting body forming a wing and having two halves that can move independently (see FIG. 17A) and also an example where only the upper half of the ejector is used in a similar manner to an air curtain but forming a throat and a diffuser with a wing flap placed on the lifting body, which matches the necessary performance (see FIG.

17B). In these examples, the flap may or may not contain primary nozzles, and the flap moves independently of the upper half of the ejector, also described as an air curtain. The advantage of such a system is that it is simpler; it still allows for high incidence angles on the wing as explained in the current disclosure avoiding wing stall through boundary layer ingestion, and the potential to independently rotate the flap and air curtain for optimized performance and maneuverability.

In particular, FIG. 17A shows examples of a planar ejector as described above with features such as dimples on a wing aft of the wing tip and such that it primarily ingests the boundary layer above the upper surface of the wing. On the other hand, FIGs. 17A through 17C show the use of an air curtain type ejector that forces ingested air (as a secondary fluid) from above the wing to accelerate and propel the aircraft forward. In all of these examples, the ejectors may rotate. Also, in another example, the ejector inlets may also rotate in a limited manner, but their diffusing walls may extend as further explained in FIG. 14, changing angles, exhaust areas, etc. as flight conditions dictate. 1401 and 201 may independently rotate about axes 102 and 202, respectively.

Fluid propulsion system and cycle

Yet another example not falling within the scope of the claims relates generally to a propulsion cycle and system providing thrust through fluid momentum transfer. The propulsion system consists of a gas generator 1) providing various streams of high pressure air or gas sources to 2) network ducts directing said compressed fluids to 3) augmenting thrust generating elements installed in the aircraft at various stations. The augmenting thrust generating elements direct a high velocity outflow jet with a velocity component in the primarily axial direction in the desired direction, thereby generating an opposing thrust force. The outflow jet is a mixture of high energy hot gases, provided to the thrust generating element through ducts from high pressure gas generator locations such as compressor bleeds, combustion bleeds, turbine bleeds and/or exhaust nozzles, and is designed to entrain surrounding air at very high entrainment velocities. The entrained air is brought into a high kinetic energy flow through momentum transfer with the high pressure gases supplied to said thrust generating element, within the thrust generating element; the resulting air-gas mixture emerges outside the thrust generating element and points mainly in the axial direction, towards said thin lifting body leading edge and mainly pressure side of the lifting body preferably in the direction to maximize lift on said downstream lifting body.

FIG. PA2-7 illustrates an example not falling within the scope of the claims, featuring the propulsion device in VTOL configuration. Ejectors 201 and 301 are oriented downward and thrust is moving the air vehicle upward. The ejectors are rotating in synchronization and the primary fluid flow is modulated to match the thrust need to the forward and rear ejectors from compressor bleeds for ejectors 201 and exhaust gases for ejectors 301.

The most efficient conventional propulsion system for medium- and long-range aircraft engines is the high-bypass turbofan. Conventional turbofans employ at least two shafts, one common to the fan and low-pressure turbine, and one common to the core, which may consist of a secondary compressor, a high-pressure compressor, and a high-pressure turbine. The high efficiency of a turbofan is driven by the high bypass ratio, low fan pressure ratios for high propulsive efficiency, and high overall pressure ratios for high thermal efficiency. The specific fuel consumption of the aircraft is inversely proportional to the product of the thermal and propulsive efficiencies. Thermal losses in a turbofan are primarily due to combustion and thermodynamic losses in components such as the compressor, turbine, and mechanical efficiencies that are less than 100%. The irreversibility of the combustion process is generally the main component leading to lower thermal efficiency and typical high pressure ratio powertrains are only 40% thermally efficient. Practicality and other aircraft limitations (weight, drag, etc.) prevent the implementation of methods known in the art for thermal efficiency improvement, such as intercooling, heat recovery, and others.

Propulsive efficiency is maximized, on the other hand, when the propeller accelerates a greater amount of air mass flow to a small axial velocity, just above and as close as possible to the aircraft's airspeed. This results in the need for very large fan diameters and high fan speeds, increasing aerodynamic drag and aircraft weight. Currently, the highest and most efficient turbofan is very large, with fan diameters exceeding 3.35 m (11 ft) in size. Although the increased fan diameter improves propulsive efficiency, drag increases due to the size of the cowling and a trade-off is usually made to obtain the ideal system. Current levels of propulsive efficiency exceed 85% and efforts are being made to distribute thrusters on wings to maximize this. A popular idea in the art is the concept of distributed thrust elements. The thrusters may be distributed on both wings and fuselages of the aircraft. Most often they are electrically or mechanically driven fans placed in wings, and receiving mechanical work or electrical energy from a central unit. Such concepts are difficult to implement due to the complexity of the network involved, the weight of electric motors and their operability at high altitudes, and in the case of mechanically driven networks, efficiency, complexity and weight. The dominant design remains the two-motor design.

A disadvantage of the current dominant design is that the turbofan is heavy and complex. Over 30% of its total weight is the fan system alone, including fan accessories and the low-pressure turbine that drives it. The large rotating parts mean that there are additional design constraints, including limitations on tip speed, restrictions on the weight and dimensions of the lower-pressure turbine, as well as inlet temperatures to the low-pressure turbine. The fan blade needs to be qualified and certified in dedicated fan blade ingestion and fan blade exit tests. In addition, the fan housing needs to contain the release of such fan blades and protect the integrity of the aircraft. With smaller systems, the challenge of scaling down a complex turbofan system is significant, if efficiency is to be maintained. Particularly for UAVs and small aircraft, bypass ratio (BPR) levels are much lower due to material limitations. As fans shrink in diameter, they need to spin faster to retain their efficiency, and tip losses occur at higher speeds leading to lower efficiency. For small turbofans, the challenge is that fan (and compressor) downscaling means that rotational speed has to increase dramatically. Those skilled in the art understand that a fan's diameter scales directly proportional to the square root of the fluid mass flow, while the fan blade tip speed is directly proportional to the product of diameter and rotational speed (e.g. Pi*Diam*RPM). Therefore, if the fan diameter is significantly reduced, then conversely the rotational speed needs to increase to preserve the same tip speed (for mechanical and compressibility reasons), otherwise performance losses increase significantly. For example, if a 50 in (127 cm) diameter fan rotates at 2000 rpm, for the same tip speed a 20 in (51 cm) fan needs to rotate at 5000 rpm, and a 10 in (25 cm) fan would rotate at 10000 rpm, etc. This also implies that the fan pressure ratio (FPR) would increase accordingly, driving lower fan efficiency in the smaller diameter range. Furthermore, containment of such a highly stressed fan component would be difficult to achieve and would incur a thicker fan casing, driving weight, and lead to significant complications with respect to the dynamics of the system rotor and its bearing subsystem. For this reason, large fans are much more efficient than smaller fans. The status quo of small turbofans is significantly lower than that of larger systems, at least 3-4 times lower than that of large BPR fans and with higher FPR, leading to low efficiencies (high fuel burn), high rotational speeds (high stress and maintenance) and difficult operability and thermal management.

Turboprops face the same challenge, although for really small systems they have the best propulsive efficiency. Their main disadvantage is the large size of the propellers required to move massive amounts of air and make their implementation difficult in systems with VTOL capabilities. The modern turboprop uses a low-pressure turbine to drive the propeller and employs additional auxiliary systems such as gears and bearings and their subsystems, pitch control, and others.

Another element of modern aircraft propulsion jets, such as turbofans and turboprops, is that a certain amount of bleed air is required outside the compressor for cabin pressurization, turbine cooling, and onboard exhaust for operability of the engine itself. The compressor bleed air of a typical modern jet engine is up to 20% of the total compressor discharge air. Compressor bleeds intended for cabin pressurization are not required if the aircraft is flying at low altitudes or is unmanned, and this portion constitutes at least 10% of the total bleed. If the turbine is not cooled, then another 10% or so of the compressor air may be removed before it reaches combustion, at the expense of lower combustion temperatures and hence cycle efficiency. However, with the advancement of new non-metallic materials and their high temperature and stress capabilities, the turbine and indeed most of the hot section can be fabricated from ceramic matrix composite materials, not only eliminating the need for cooling air but also allowing for higher firing temperatures. For example, while the turbine inlet temperature limitation with current uncooled metallic components is known to experts familiar with the subject to be approximately 954 °C (1750 °F), current CMC materials could withstand an uncooled turbine firing temperature of 1093 °C (2000 °F) or higher. This results in much higher cycle efficiency and in most cases reduced engine weight, with overall benefits to the aircraft. If a 954°C (1750°F) combustion temperature, uncooled, all-metal component engine with 20% compressor bleed air is replaced by a 50% bleed air compressor burning at 1093°C (2000°F) with ceramic components, then the cycle efficiency can be comparable, while 50% of the compressed air is made available for other purposes at the compressor unload station.

Table 1 shows such a comparison for various air bleeds for two uncooled engines with the same unit flow (i.e. 1 kg/s) and various bleed percentages and the same power output of the turbine supplying the required input power to the compressor. The first line shows the cycle pressure ratio, the second line shows the compressor bleed, the calculation of the thermal efficiency of the metallic engine is shown in line 3 and the thermal efficiency of the CMC versions with similar bleeds and the maximum bleed at the same efficiency compared to the metallic version are shown in the last two lines. The general assumption is that the turbine is not cooled in all cases, but air is bled out of the compressor for another purpose. The table shows that if the bleed percentage is kept the same between the metallic and CMC versions, then at a cycle pressure ratio above 8, the CMC engine becomes more efficient. On the contrary, if the engine efficiency is to be kept similar to the metallic one, the bleed air percentage can be drastically increased. This can also be explained by maintaining the same fuel flow to the combustion but reducing the air flow until the combustion temperature becomes 1093 °C (2000 °F) (uncooled CMC technology) from 954 °C (1750 °F) (uncooled maximum metallic technology). By producing the same power output from the turbine to balance the power input of the compressor, more compressor air can be made available for higher ignition temperatures. On this basis, the inventor has devised a cycle which allows a large amount of compressor bleed to be routed through a flow network and supply a series of distributed augmentation thrust devices placed at locations which increase and improve the propulsive efficiency of the current state of the art, to similar or better thermal and overall efficiency of the aircraft.

Table 1

Consequently, conventional thrusters cannot be scaled down without significant compromise to their efficiencies. An example not within the scope of the claims overcomes the current drawback through the use of an improved cycle that eliminates the fan subsystem along with the low-pressure turbine. As such, this is a particularly suitable system for smaller aircraft and UAV systems, particularly those that need to be capable of VTOL and STOL operation, due to its efficient, compact and highly integrated powertrain systems.

The propellant consists of a "chopped" fan and a high-pressure compressor placed on the same shaft with a high-pressure turbine to form a gas generator, a duct network connected to ejectors, and the thrust boosting ejectors. The cycle utilizes a compressor system consisting of a chopped fan (core precompression only) and a single or multistage high-pressure compressor, preferably a centrifugal compressor with multiple bleed ports. The compressor bleed ports can bleed up to 50% of the total airflow in the system, with the remainder directed to a combustion chamber system. Combustion adds heat in the form of fuel at constant pressure or volume, and generates a hot gas stream that is directed to a turbine. The high-pressure turbine expands the hot fluid to a pressure and temperature lower than the turbine inlet pressure and temperature in a conventional expansion process. Preferably, the turbine and combustion are high-temperature materials requiring little or no cooling flow, such as modern CMCs. The turbine, which may be centripetal or axial in at least one stage, supplies the work required to drive the compression system. The exhaust gas exiting the turbine is at pressures and temperatures lower than the turbine inlet conditions, but at least twice the ambient air pressure and at temperatures typical of today's turbofan low-pressure turbine levels, i.e., 815–982 °C (1500–1800 °F). Thus, the high-pressure turbine expansion process still results in a high-energy, high-temperature, high-pressure flow that instead of being directed to a low-pressure turbine, is directed through ducts to various locations on the aircraft to a fluidic thrust generating propellant.

The ducts may also be insulated and use high temperature materials such as CMC. The propellant elements receiving the pressurized air or hot gases are employing fluids for entrainment and acceleration of ambient air in a first section; and after mixing the motive fluid and ambient air and completing a complete momentum transfer from high energy to low energy (ambient air), accelerating the mixed flow in a second diffusion section, thereby supplying a high velocity jet outflow as a mixture of the high pressure gas (supplied to the propellant from a gas generator) and ambient air entrained at high velocities, preferably and primarily in axial direction, with a certain axial velocity profile known in the art. The entrainment velocities of such propellant elements are between 3-15 and up to 25 per part of high pressure fluid supplied. Due to the high entrainment and turbulent and complete mixing of the flows, the jet outflow is consequently much lower in temperature. Following the physical law regarding mixing and momentum transfer, the velocity of the jet outflow from the propellant element is close to but exceeding the airspeed of the aircraft. The jet outflow is also non-circular in nature, with little or no rotational components (as opposed to the large propellers of a turboprop or even a turbofan) and may be directed to the lifting bodies to further recover some of their energy after producing thrust, for example directed towards the leading edge of a short wing placed some distance behind the propeller to generate additional lift. In all examples the gas generator is a modified turbofan in which the fan has been cut off to provide only core flow.

FIG. 26A shows a traditional bypass turbofan that bypasses most of the flow and mixes the core and bypass flows at the turbofan exit. FIG. 26B shows the turbofan with a cut-out fan allowing core only and bleed flow out of the compressor to produce the thrust necessary for propulsion. The compressor bleeds are conveniently located across the cold section 2601 of the gas generator 800, as opposed to a hot section 2602 of the generator, so as to allow for maximum efficiency at any time during operation. For example, at takeoff, more compressor bleed may be required and higher rotor speed may optionally be advantageous. At this portion of the mission, the bleed ports are opened so that compressor operation is away from the surge line at a more advantageous condition than if no bleed were present. A current turbofan such as that presented in FIG. 26A may allow only a maximum of 15% bleed throughout the mission, but by changing the slit fan design in favorable ways, more core flow and more compressor bleed can be induced in the present disclosure, up to and including 50% of the total bleed air through the engine. Multiple bleeds may also be involved, to improve system efficiency, maximize bleed at lower stages, and minimize bleeds at higher pressure, as will be recognized by those skilled in the art. However, the amount of flow intentionally involved in the bleed cycle alone is particularly applicable in this invention. Bleed air from the compressor bleed ports is directed through ducts to booster thrust thrusters positioned along the upper surface of a lifting body or immediately aft of a lifting body at a high angle of incidence.

FIG. 27A illustrates an example of a purge and conduit network embodied in the present invention. The network contains a gas generator 800 that feeds a plurality of cold boost ejectors 801 and hot boost ejectors 901 through compressor purge ports 251 and 351, respectively. Pressure and temperature sensors may take flow measurements as signals from elements 1702 (cold) and 1707 (hot) are fed to microcontroller 900 (not shown). Flow from gas generator 800 to thrust boosting ejectors 801 via compressor bleed ports 251 is dictated by control valves 1703 controlled by microcontroller 900. The same controller dictates actuation of rotary joints 1701 (for element 801) and 1705 (for element 901). FIG. 27A further shows an array of four (4) cold thrust boosting ejectors 801 fed from the same compressor ports 251 of said gas generator 800 and controlled by microcontroller 900.

FIG. 27B depicts a network similar to that illustrated in FIG. 27A, but with only two cold thrust boost ejectors that are fed by compressor bleed ports outside of gas generator 800. Rotating joints 1701 allow rotation of elements 801 in multiple directions and may also allow passage of fluid to said ejector 801. The position of said ejector relative to the aircraft is controlled through electrical or pneumatic or mechanical means from a microcontroller 900. Sensors 1702 for measuring flow rates, pressures and temperatures in the conduit downstream of bleed port 251 are used to feed information to the microcontroller. In turn, the microcontroller commands rotation of elements 801 about rotating joints 1701 while commanding flow adjustment through control valves 1703. Similarly, the thrust magnitude and orientation of elements 901 may be adjusted by flow rate adjustments through control valves 1707 and orientation adjustments about rotating joints 1705 until the aircraft attitude is acceptable. The controller is thus fed information from the propulsion system duct network, as well as the operating parameters of the gas generator and the thrust magnitude and orientation at each of the thrust augmentation ejectors.

FIG. 27C shows the microcontroller 900 and its network, which represents at least twelve (12) inputs and at least four (4) outputs. The outputs primarily control the flow rate and thrust (ejector) orientation to control the attitude of the aircraft at any time in its mission.

FIG. 27D provides further details of the network. Flows from compressor bleed ports 251 and exhaust 351 are fed to thrust boost ejectors 801 and 901. Inputs to microcontroller 900 include gas generator parameters (rpm, compressor bleed air temperature and pressure, exhaust pressure and temperature, etc.) via input 11. Inputs 26 include feeds from accelerometers included in the system. Input 30 is the gyroscope. Input 40 is an ultrasonic or barometric altitude sensor input that signals the altitude of the aircraft. Input 50 is the GPS input. Inputs 70 are the Bluetooth input. Inputs 80 are the R/C receiver.

Furthermore, FIG. 27D illustrates feedback from sensors in the conduits, shown as feed information 2702 and 2706, to the controller for flow adjustments through actuation of control valves 1703 and 1707. The control valves are connected to the controller via cables 2703 and 2707 respectively, and are adjusted based on input received from the controller. Flow is adjusted and measured accordingly by sensors 1702 and 1706, which provide feedback information to the controller to signal that the adjustment was made accordingly. Similarly, the other sensors 11, 26, 30, 40, 50, 60, 70, and 80, either individually or together, are processed by the controller and adjustments to the positions of ejectors 801 and 901 are transmitted via cables 2701 and 2705.

In another example not within the scope of the claims, the propeller may rotate downward to direct thrust to change the attitude of the aircraft for vertical takeoff or short takeoff. FIGS. 28A-28E show possible forms of a propeller in the present disclosure.

FIG. 28A illustrates a relatively simple four-ejector system with the ejectors being fed by a gas generator in the center. Two of the (cold) ejectors are fed from the compressor bleed ports with compressed air as the motive (or primary) air, while the hot ejectors receive exhaust gas from the gas generator exhaust port. All four ejectors point downward, in hover mode, but aircraft attitude adjustments are possible through changes in the parameters mentioned in FIGs.

27A - 27D. FIG. 28B shows an example not falling within the scope of the claims where the device(s) may be integrated into an aircraft. In FIG. 28B, only the two cold ejectors (of four) are shown. The ejector is flat, and is positioned aft of the main wing apex, working to expand the main wing stall margin and generate high speed outflow for use on downstream lift generating structures. The flat ejector may also rotate along the wing main axis to hover (pointing downward toward the ground, primarily) or to adjust altitude in flight. FIG. 28C illustrates a more complex canard-type system consisting of four ejectors (two hot, located in the tail, and two cold, positioned aft of the canard lift bodies) and the ejector embedded within the canard ejector-side (cold) housing system. Ejectors are flat and produce a rectangular shaped jet outflow in their exit plane, primarily in the axial direction of flight. Alternatively, an ejector having 3D elements is shown in FIG. 28D, where the inlet and throat and diffuser are 3D in nature (rather than 2D), improving drag and overall performance. In FIG. 28E, only an upper half of a flat ejector is depicted above the wing, mating with the wing flap to form a complete structure with the primary fluid only introduced into the upper half of the ejector (half-ejector, mating with the wing flap).

FIG. 29 shows a possible arrangement of the propulsion system in takeoff or hovering flight in an example not falling within the scope of the claims. The ejectors point downward to raise the fuselage and maintain it in a hovering position.

FIG. 13 shows the maneuverability of a UAV equipped with a propulsion system. The pitch, roll, and yaw positioning of the ejectors is shown for both tail (hot) and cold (canard) ejectors.

In one example not within the scope of the claims, the propellers receiving both compressor discharge air (at least twice ambient air pressure) and hot gas efflux from the high pressure turbine (at least twice ambient pressure) are in this example pointed downward at takeoff, thereby generating counter thrust that overcomes the weight of the UAV for takeoff. U.S. Patent No. 8,087,618 (Shmilovich et al.) discloses the use of such a device integrated into the wing system and utilizing a jet engine exhaust to direct exhaust air or compressed air only into takeoff and lesser portions of compressor bleed air (less than 15% is mentioned) for additional flow control over a wing. In particular, they do not increase thrust but simply redirect the exhaust stream by controlling it with compressed air during takeoff. An example not within the scope of the claims utilizes a specially designed powertrain system that particularly draws bleed air, in excess of 20%, from the compressor and directs it to said propellant throughout the flight, from takeoff to landing. The particular way of achieving this is by designing a compressor with more open first stages that can accommodate more flow, followed by purging a portion of the flow in large quantities, such as up to 50% of the total air flow, and directing this portion at all times to the cold gas propellants, and using the entire remaining portion of the flow for the thermodynamic cycle with the residual energy of the flow after the high pressure turbine directed to a hot gas propellant. Compressor bleed flows can also be modulated by employing flow controls, such as control valves or fluidic valves that modulate the flow delivery to the propellants. Both types of cold and hot thrusters can rotate at least 90 degrees to 120 degrees pointing down and up compared to the forward flight direction and independently. The cold gas thruster can be embedded with, or hidden in, the wings, or preferably in the wake of a first very high incidence angle wing (canard wing) and improve its stall margin significantly by placing the thruster inlet in proximity to the canard lifting body and preferably in the last third of its chord and closer to the trailing edge. High incidence angle can cause separation and stall, but the addition of the thruster at such a location will extend its operability well beyond the stall point.

In another example, the injection of fluids such as water or liquid nitrogen to cool the hot gases supplied to the hot propellant can increase the takeoff thrust generated by said propellant through the increase in the mass flow rate of the driving air. In case the propulsion system is embedded in a UAV, the amount of water on board the aircraft can be such that after takeoff and the end of the mission, when the onboard fuel has almost been consumed, the landing will not need the additional thrust by at least 25% and up to 50%.

In yet another example, the use of high pressure turbine exhaust gas as the primary/motive fluid to the hot ejector may be augmented with additional cold air compressor bleed, particularly to maintain a cooler mixture feed temperature to the main hot propellant primary nozzles, particularly during horizontal flight. In this manner, with constant pressure mixing and decreased gas mixture temperature, longer life and/or cheaper duct materials may be utilized. Modulation of the cold compressor bleed air may be accomplished through a valve switching flow from the cold propellant supply to the duct feeding the hot propellant, or through a secondary inlet to the hot propellant plenum to extend its useful life. In this case, the cold propellants are either inline with the main wing system or may be retracted within the fuselage and thus not participate in thrust generation.

The thermodynamic cycle of a typical jet engine is presented in FIG. 30A. The evolution of the working fluid is described from the end of the inlet (point 2) to 3 through a compression process, fuel addition and constant pressure combustion through a 3-4 isobaric process, expansion over a 4-5 turbine. The latter provides the work required by the compressor and the additional energy available to drive a fan (via a turbine connected to the fan) for a turbofan or for direct expansion through a nozzle to the atmosphere, in the case of a turbojet. This example eliminates the free turbine required to drive the fan, connects the chopped fan to the engine mainshaft and uses the energy of the combustion gases at point 45 to drag and boost thrust through a specially designed ejector incorporated with the aircraft, at all points during the aircraft mission (takeoff, transition, flight level, hover and landing). The thermodynamic evolution of the invented cycle takes the gases through a nearly isentropic expansion at lower pressure at high efficiency, much higher than the expansion through a turbine. The process 45-A' describes the evolution and can be recognized as nearly isentropic since such nozzle expansions are known to have very high efficiency. The evolution of the working fluid 45-A' occurs through a multitude of primary slots located within the ejectors described above. The expansion is continued by constant pressure or constant area mixing with ambient air approaching p<2>static conditions. The evolution follows A'-D' for the working fluid while ambient air is brought at constant pressure from the inlet condition point C to D'. In this constant pressure mixing process, the final temperature of the mixture depends on the ambient air entrainment ratio in the ejector. As described below, a specially designed ejector is utilized in this invention which maximizes the entrainment ratio across various elements of said primary nozzles and the mixing section of the ejector to values beyond 5:1 (five parts entrainment air to one part primary working fluid). A pumping effect then occurs, raising the temperature and pressure of the mixture of warmer primary fluid and entrained ambient air to Tmix and Pmix respectively, in excess of pamb. This is point D in the diagram of FIG. 30B. A near isentropic diffusion and ejection of the mixture is the D-E evolution, with final temperature and pressure of Tout and Pout respectively, where Pout equals ambient pressure at aircraft speed. The location of point D is between point 2 and point 5, closer to point 2 for higher entrainment ratios. The advantage of such a system is then apparent, seeing that a large amount of air can be entrained and energized to produce thrust at lower temperature and mixture velocity. This, in turn, allows the exhaust from such a thermodynamic cycle to be utilized not only for thrust generation but also to be advantageously directed over various lifting bodies for additional lift generation, or vectored for VTOL and STOL capabilities of the aircraft, at the exit of the cycle. Furthermore, the placement of the entrainment inlet of such ejectors in some examples may be such that ingestion of the resulting boundary layer of a lifting body such as a wing is obtained, with additional benefits to the wing stall margin of the assisting lifting body. In one example, a first set of wing lifting bodies is positioned such that it operates at very high angle of incidence, with very little margin for stalling. By placing said ejector just behind said lifting body apex point, in the area prone to developing boundary layer separation, the suction side (i.e. the drag side or inlet side of the ejector) determines that the stall margin is significantly improved, allowing very high lift generation to be obtained on said lifting body/wing, free of stall.

Also, in some cases where a short takeoff distance is desired, the exhaust gas from a turbofan can be directed onto the suction side of a lifting body (such as a fin). Although several concepts have used this technique, there have been limited results. In this example, there is higher lift proportional to the higher local velocity squared, at least for the portion of the wing exposed to the emerging jets of the thrust element, because it utilizes the benefits of directing the higher kinetic energy fluid (exhaust gas mixture air) directly to the pressure side of the wing or fin, rather than the suction side, or directly to the leading edge much in the manner of an lifting body of turbo machinery (such as a turbine).

Furthermore, the ejector exhaust which is at a significantly lower temperature and still at a higher average exit velocity than the airspeed, can be directed towards a secondary, downstream, thin lifting body. The ejectors direct their efflux jets towards lifting bodies which are thin and can be made of reinforced composite materials. The higher efflux jet velocity determines a high lift force on such lifting body compared to the lift generated by such lifting body would then receive only the airspeed flow of the aircraft. In contrast, the size and shape of the lifting bodies can be significantly reduced to produce a lift force similar to that of a very large wing. Returning now to the inlet of the ejectors and observing their placement just above the lifting body 803, the ingestion of the boundary layer developed beyond the apex of said lifting body 803 into the ejector and this boundary layer being drawn in by said ejector determines a better flow stall margin of the lifting body 803 and allows it to operate efficiently at a higher angle of incidence.

In another example, not falling within the scope of the claims, shown in FIG. 31, said cold ejectors 801 are placed behind the lifting body 803 and in front of the lifting bodies 802, still impacting both increasing the stall margin on the 803 due to the boundary layer suction of said lifting body 803 and the lift force due to the higher velocity of the efflux jet at the exit of 801 efficiently directed towards the lifting body 802. This allows for a more aggressive horizontal flight attitude of both lifting bodies 803 and 802, shorter lifting bodies for the same lift and ejector twist for vertical takeoff, hovering and maneuvering of the aircraft. It also allows for a more beneficial use of the efflux jet from the thermodynamic device to be used towards lift generation and not wasted to the environment, as is the case for current state of the art jet engines.

Although this is not intended to be an exhaustive list, different embodiments of the present invention are designed to provide some or all of the following improvements and advantages:

improve the ability to maximize thrust augmentation and vector a Coanda-type flat ejector jet outflow under all flight conditions;

improve efficiency and shorten the device for better integration with the aircraft wing or fuselage by introducing particular 3D features into the primary nozzle and Coanda surface;

embedding such a device with a wing to take advantage of the particular geometries of the wing in order to improve the efficiency of the aircraft;

improve the efficiency of the primary nozzle to entrain secondary fluid and mix in the shortest period and length of the device through additional features;

improve the overall geometry in a non-circular manner to allow for efficient operation in horizontal flight of the aircraft, as well as takeoff, hover and landing, while improving the propulsive efficiency of the aircraft and eliminating the presence of nacelles and main propulsion engines on the wings and fuselage of the aircraft;

generate additional thrust and lift due to increased local jet velocity over the wing using the residual kinetic energy of a jet outflow that normally only generates thrust through a mechanical connection;

shortening the wings while preserving the same lift through the extension of the propeller diffuser walls as a propulsion and lift generating device;

improve the ejector to work better in conditions far from the ideal conditions of a fixed geometry ejector (e.g. optimizing operations and thermodynamic propulsion cycles by using 2 halves of an ejector, being able to move them relative to each other, and adding a fin-like feature to fully expand and collapse the walls of the ejector diffuser);

increase the lift-to-span ratio due to the relatively low temperature of the jet outflow mixtures exiting the propellant and an axial velocity component at values higher than the aircraft speed;

include composites as a type of material used in thin airfoil due to their ability to withstand higher temperatures of emerging mixture jet outflow;

reduce the overall dimensions and weight of the aircraft because the lifting body can be thinner in width and shorter in wingspan with high mechanical resistance to stress;

significantly improve the maneuverability and versatility of an aircraft, including enabling V/STOL and hovering flight, through oscillation and modulation of both the propellant and lifting body flow; and/or

improve aircraft attitude control, hovering and VTOL capabilities by enabling a compact system with small roll and a distributed propulsion system, particularly in UAVs, UAS and drones.

Furthermore, in addition to the many features mentioned above, different embodiments of the present invention may also have some or all of the following improvements and advantages:

The thermodynamic cycle is simpler, with an ejector/eductor type element replacing the entire functionality of the low pressure fan and turbine subsystem, thereby reducing the weight of the system by at least 30%. This is particularly advantageous for smaller UAV type systems where turbofans are not effective due to the reasons explained above;

the potential to spin or vector eductor-type thrusters independently and allow vertical takeoff and landing, without moving large rotating parts;

the potential to modulate flow to these thrusters during takeoff and level flight, as well as landing and emergencies, thereby applying different levels of thrust at various locations on the aircraft, and to completely isolate any number of such thrusters;

the potential to eliminate large rotating part components with non-moving parts of the same functionality, i.e. fan replaced by fluidic thruster/eductor; a direct improvement in component lifetime of non-moving versus rotating parts is expected, especially for small UAVs and aircraft where fan dimensions require very high speeds;

the potential to use lightweight, high-temperature materials such as composites, carbon fiber-based materials and CMC for ducts and propellants;

the potential to modulate purges so that in horizontal flight only the hot propellants are supplied with hot gas or a mixture of hot exhaust gas and cooler compressor air is purged from the gas generator;

the benefit of the gas generator being operated at the same rotational speed optionally advantageously without large deviations in RPM between takeoff and cruise, away from the overload or stall line;

the benefit of giving any shape to the propeller and being able to integrate it to a great extent with the fuselage and wings of the aircraft;

the benefit of having high drag and turbulent mixing within such propellants such that the jet outflow from their exhaust is sufficiently low in temperature to allow a lifting body for aircraft attitude or lift control to survive and function properly, including generating more lift using the higher velocity jet; and/or

the benefit of embedding the propellant in the wing behind the wing roll apex where the boundary layer would otherwise separate at high incidence angles, thereby ingesting the boundary layer and delaying its separation and increasing the high margin of the wing in level flight.

It should be noted that any of the ejectors 701, 801, 901 may be configured using any ejector geometry described herein.

Although the foregoing text sets forth a detailed description of numerous different embodiments, it should be understood that the scope of protection is defined by the words of the claims that follow. The detailed description should be construed as exemplary only and does not describe every possible embodiment because describing every possible embodiment would be impractical, if not impossible. Numerous alternative embodiments could be implemented so long as they remain within the scope of the claims.

Therefore, many modifications and variations can be made in the techniques and structures described and illustrated herein without departing from the scope of the present claims. Accordingly, it is to be understood that the methods and apparatus described herein are illustrative only and do not limit the scope of the claims.

Claims (10)

1. A vehicle (804), comprising: a main body having a bow part, a tail part, a starboard side and a port side; a gas generator (800) coupled to the main body and producing a gas stream; at least one bow conduit coupled for fluid transmission to the gas generator (800) for receiving gas from the gas stream; at least one tail conduit coupled for fluid transmission to the gas generator (800) for receiving gas from the gas stream; first and second bow ejectors (801) fluidly coupled to the at least one bow conduit for receiving the gas from the at least one bow conduit, coupled to the bow portion and coupled respectively to the starboard side and the port side, the bow ejectors (801) respectively comprising an outlet structure out of which the gas flows from the at least one bow conduit at a predetermined adjustable rate; at least one tail ejector (901) fluidly coupled to the at least one tail conduit for receiving gas from the at least one tail conduit and coupled to the tail portion, the at least one tail ejector (901) comprising an outlet structure out of which gas from the at least one tail conduit flows at a predetermined adjustable rate; first and second primary lifting body elements (802) having leading edges, the primary lifting body elements (802) being coupled respectively to the starboard side and the port side, the leading edges of the first and second primary lifting body elements (802) being respectively located directly downstream of the outlet structure of the first and second bow ejectors (801) such that gas from the bow ejectors (801) flows over the leading edges of the primary lifting body elements (802); and at least one secondary lifting body element (904) having a leading edge and coupled to the main body, the leading edge of the at least one secondary lifting body element (904) being located directly downstream of the outlet structure of the at least one tail ejector (901) such that gas from the at least one tail ejector (901) flows over the leading edge of the at least one secondary lifting body (904); 1. A vehicle having a canard wing (803) coupled to the forward portion and coupled to the starboard side and the port side respectively, the canard wings (803) being configured to develop boundary layers of ambient air flowing over the canard wings (803) when the vehicle (804) is in motion, the canard wings (803) being respectively located directly upstream of the first and second forward ejectors (801) such that the first and second forward ejectors (801) are coupled for fluid transmission to the boundary layers, wherein the first and second forward ejectors (801) respectively comprise first and second inlet portions for intake of upstream air, and the first and second forward ejectors (801) are positioned such that boundary layers developed by the canard wings are ingested by the inlet portions. 2. The vehicle (804) of claim 1, wherein the gas stream produced by the gas generator (800) is the sole means of propulsion of the vehicle (804). 3. The vehicle (804) of claim 1, wherein: The gas generator (800) comprises a first region in which the gas stream is at a low temperature and a second region in which the gas stream is at a high temperature; the at least one bow duct provides gas from the first region to the first and second bow ejectors (801); and The at least one tail duct provides gas from the second region to the at least one tail ejector (901). 4. The vehicle (804) of claim 1, wherein the gas generator (800) is arranged in the main body. 5. The vehicle (804) of claim 1, wherein the first and second bow ejectors (801) each have a leading edge, and the entirety of each of the first and second bow ejectors (801) is rotatable about an axis oriented parallel to the leading edge. 6. The vehicle (804) of claim 1, wherein the first and second bow ejectors (801) each have a leading edge, and the entirety of each of the first and second bow ejectors (801) is rotatable about an axis oriented perpendicular to the leading edge. 7. The vehicle (804) of claim 1, wherein the at least one tail ejector (901) has a leading edge, and the entirety of the at least one tail ejector (901) is rotatable about an axis oriented parallel to the leading edge. 8. The vehicle (804) of claim 1, wherein the at least one tail ejector (901) has a leading edge, and the entirety of the at least one tail ejector (901) is rotatable about an axis oriented perpendicular to the leading edge. 9. The vehicle (804) of claim 1, wherein at least one of the output structures is non-circular. 10. The vehicle (804) of claim 1, further comprising a cabin portion (805) configured to allow manned operation of the vehicle (804).